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Aerospace Engineering is a branch of engineering that focuses on the design, development, testing, and production of aircraft and spacecraft. It encompasses various disciplines, including aerodynamics, materials science, structural analysis, propulsion, and avionics, to ensure the safe and efficient operation of vehicles in the atmosphere and outer space.
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CONTOH PROPOSAL PKM-KC (PROGRAM KREATIVITAS MAHASISWA KARSA CIPTA) (Didanai DIKTI 2018)
by Meda Aji Saputro
Proposal ini sudah di publikasikan pada simposium RAPI 2018
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Manufacturing Technology for Aerospace Structural Materials - F.C. Campbell
by Francisco Dávalos
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Engineering Mechanics Statics 13th Edition-Solution Manual
by Ozzie Sahan
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Automatic Flight Control System: Classical approach and modern control perspective
by Agus Budiyono
The document was used as a lecture note for the graduate course on flight control system at the Malaysian Institute of Aviation Technology (MIAT), Kuala Lumpur, Malaysia. Some parts of the document represent an enhanced version of the... more
Classical approach and modern control perspective Automatic Flight Control System
Figure 1.1: Autopilot Control System Diagram: Example of Pitch Channel This autopilot system does not work continuously but only at a certain pe- riod of time. The AP system can be activated by turning the AP switch and deactivated by over-riding the system through the movement of the control ma- nipulator. Therefore the characteristic of the AP system is of limited authority. Fig. 1.1 shows the functional diagram of the autopilot system used in the pitch longitudinal channel of an aircraft.
Figure 1.2: Stability Augmentation System Diagram. Example of Pitch Channel The type of automatic flight control system that adds stability to the aircraft is usually called the Stability Augmentation System or SAS. This type of auto- matic flight control system improves the stability of an aircraft at certain flight configurations and conditions within the flight envelope. For conventional air- crafts, the stability augmentation will be needed during the flight at low speed and low altitude for instance during landing or approach. The control opti- mization of typical aircrafts are conducted only at a certain flight configuration such as cruise configuration. This makes the aircraft stability at other flight configurations namely approach, landing or other special configurations tend to deteriorate. The stability augmentation is therefore necessary for those configu- rations. The stability augmentation can be achieved by increasing the damping ratio of the existing aerodynamic damping ratio through the application of feed- back control system.
Figure 1.3: Command Augmentation System Diagram. Example of Pitch Chan- nel
Figure 1.4: The Stability Provider and Control Power Optimizer Control System Diagram
Figure 1.6: The aicraft cockpit as the interface between controller and controlled systems
Another very critical element of the controller is the information processor sys- tem. In the manual controller (human controller), this system is represented by the pilot’s brain supported by the basic information processing computers to speed up the decision making process. In the automatic controller, the infor- mation processing element is represented by a Flight Control Computer (FCC). The FCC works continuously in real time depending on the authority level of the implemented automatic control system. The software inside the FCC that manipulates the input of the FCC to be converted to the desired control signal by the control system is called the control law. The control law can be in the form of simple instructions which typically used by the autopilot. Some of the examples of control law are:
Figure 1.7: Mechanical Control System: example of longitudinal channel
Figure 1.8: Hydraulically power assisted control system: example of longitudinal channel
Figure 1.9: Hydromechanical Control System
Figure 1.10: Electrohydromechanical Control System
Figure 1.11: Electrohydraulic Control System
Figure 2.1: SAS as an inner loop of aircraft autopilot system
Figure 2.4: The location of auto pilot system components in the standard control diagram variable selected as the reference. On condition that the AP system is working, the pilot can release his hand and feet from the control manipulators, wheels, and controller stick or pedal. All he needs to do is occasionaly checking up on the aircraft trim condition as displayed by the Attitude-Direction-Indicator (ADI). To disengage the AP system, the pilot would just need to grasp controller stick/wheel/pedal and to displace it a little bit. In this case, the AP control loop will be automatically disengaged and the pilot regains the full control of the aircraft. With this characteristic, the autopilot is often referred to as a low authority system.
Figure 2.5: Functional diagram of pitch attitude hold system
with the static sensitivity coefficient of The longitudinal mode transfer function of the N250-100 aircraft during cruise at the altitude of h = 15000 ft and the speed of V, = 250 kts is given as the following (for the elevator control channel): As a case study, the pitch attitude hold system of the N250-100 aircraft— prototype 2 Krincing Wesi— manufactured by the Indonesian Aerospace Inc. is presented. See Fig.2.7. In the case of a conventional aircraft, Goz(s) will have two real zeros and five poles consisting of two pairs of complex poles of the aircraft and one real pole associated with the autopilot control mechanism servo. Observing the closed loop characteristic polynomial, Eq.2.7, note that K., (servo gain) is the only gain that can still be changed or varied. The vertical gyro sensitivity S,g and servo lag factor — are typically of a constant value.
Figure 2.8: Root locus of Pitch attitude hold system for N250-100 PA-2 aircraft
Figure 2.10: Time response of 6(t) due to step 0,¢f = 5° with and without pitch attitude hold system. It is apparent that for the case with the pitch attitude hold, the value o! 9 can be quickly maintained in less than 15 time units, while when the pitch attitude hold is not present, Kgs, = 0, the time it takes to reach the steady leve of Ore y = 0 is more than 500 time unit or 35 times as along as that of the syste with the pitch attitude hold. Also note that when the pitch attitude hold is off. there exists an offset angle of about 15° whereas with the pitch attitude hold the offset angle is only approximately 0.25°. Fig.2.11 shows the time respons« of velocity, u(t), and angle of attack, a(t). The bleed-off speed to hold the pitck angle @ = 5° is about —35 kts while the angle of attack increases by 0.5°.
Figure 2.11: Time response for u(t) and a(t) for K.~ = —8.969!
Figure 2.12: Speed hold system functional diagram the value of the reference speed te ¢(t). In this condition the signal of speed difference €,, from the computer will be zero. When the autopilot is working, the pilot can release the power lever and let the engine control manipulator moves by itself following the closed loop process of the speed hold autopilot system. Hence this type of flight speed holding system is commonly called an Auto Throttle system.
Figure 2.13: Mathematical diagram of speed hold system
Substituting Eqs.2.24b—2.28 to Eqs.2.22—2.23a, the following equation is ob- tained The following figure shows the root locus diagram of the speed hold system of the N250-100 aircraft for positive K,,5, gain. The root locus shown in the figure is varied as a function of the gain K = KpcK-. Notice that the pitch oscillation mode does not change at all since it is constrained by the two complex zero of Ns._.(s) = 0. Nonetheless, the damping ratio of the phugoid mode increases up to its maximum value before moving to the unstable region.
Figure 2.15: Root locus diagram for the speed hold system of N250-100 PA-2- zoomed around the pitch oscillation and phugoid modes
Figure 2.16: Time response of u(t) to maintain u,-rs = 1 with and without the speed hold system, K* = 17.3838
Figure 2.17: Time response of a(t) and @(t) to maintain u;-er = 1, with the speed hold gain K* = 17.3838
The equation of the rate of change of flight altitude can then be obtained
Figure 2.19: Outer-wing air brake (Airbus A-320/319/321) In the Laplace domain, the flight altitude can then be obtained as:
Figure 2.20: Functional diagram of altitude hold system
Figure 2.21: Kinematic diagram of aircraft rate of climb A(s) hres (8) Based on Fig.2.22, the closed loop transfer function can be written as follows:
Figure 2.22: Mathematical diagram of the altitude hold system and the denumerator polynomial is:
Figure 2.23: Root locus for the altitude hold system for the N250-100 with h —> 6 feedback with gain Ky <0
Figure 2.24: Root locus for the altitude hold system for the N250-100 with h — 6, feedback with gain K.. < 0-—zoomed to show the phugoid mode Note that Sp is the conversion gain from signal €), to signal @;¢¢ which is a constant. From the elaboration given in Sec.2.3.1, a good pitch attitude hold control loop has been obtained with Kz, = —8.9695. The corresponding roots of the closed loop polynomial have been given in Eq.2.15. Thus the inner closed loop characteristics polynomial can be expressed as:
Figure 2.25: Altitude hold system with an attitude hold system as the inner loop From the above diagram, the equation for h(s) can be expressed as:
Figure 2.26: Altitude hold system with inner loop having gain Kk? If the gain is chosen such that: the following characteristic roots are obtained:
Figure 2.27: Root locus diagram of the outer loop of the altitude hold system for N250-100 PA-2 aircraft
Figure 2.28: Root locus diagram of the outer loop of the altitude hold system for N250-100 PA-2 aircraft-zommed to show the phugoid mode Using the model of control mechanism and elevator servo as given in Eq.2.4 with the value of +, = 0.1 sec., the polynomial of the inner control loop using the forward acceleration feedback can be expressed as follows:
Figure 2.29: Time response of h(t) for an input of h,;-, = 1 from a system with and without altitude hold for N250-100
Figure 2.30: Altitude hold system with forward acceleration a, as the inner loop feedback The outer closed loop transfer function relating h(s) and hres(s) can then be derived as follows:
Figure 2.31: Altitude hold system with inner loop using forward acceleration feedback az
Figure 2.32: Root locus diagram of inner control loop 1 —> 6¢ f altitude hold system
Figure 2.33: Root locus diagram of inner control loop u — 6, for N250-1 altitude hold system—zoomed around the phugoid mode
Figure 2.34: Root locus diagram of outer control loop % — 6. for N250-1( altitude hold system
Figure 2.35: Time response of h(t) for N250-100 altitude hold system with the inner loop of a, —> 6, feedback 3. Comparator in the autopilot computer transfer function:
Figure 2.36: Functional diagram of Bank angle hold system namely fly by wire or fly by light. The block of the AP servo and control mechanism can then be modeled as a system with time response of T, representing the first order lag factor as follows:
For a conventional aircraft, Gox(s) typically has a pair of complex zeros and five poles consisting of a pair of complex poles and three real poles where one of the poles is associated with the control mechanism and servo system.
To have a more elaborate understanding of the work principle of the bank angle hold system, a case study will be taken for the Indonesian Aerospace N250- 100 PA2 ”Krincing Wesi” at the cruise flight configuration with a low speed of V = 180 KEAS and at the altitude of h = 15000 ft. The lateral directional transfer function for aileron input for the aircraft is given as follows:
Figure 2.38: Root locus diagram of the bank hold system of N250-100 aircraft for a number of 7, values
Figure 2.39: Time response of y(t) for a = 1/7, = 10,5,2 due to an impulse function input
Figure 2.40: Time response of y(t) for a = 1/7, = 10,5, 2 due to a step function input
Figure 2.41: Forces equilibrium during the turn maneuver
Figure 2.42: Functional diagram of heading hold system specified by appropriately selecting the gain ks Sa;, Hence the transfer functior of the inner loop can be expressed as:
Figure 2.43: Mathematical diagram of heading hold system
Figure 2.44: Heading hold system with bank angle hold as an inner loop In this case, the characteristic polynomial for each value of a can be given as the following:
Figure 2.45: Root locus diagram of the heading hold system of N250-100 aircraft for a number of a = 1/7, values
The time response is presented in Fig.2.46. It is clearly indicated that the smaller the value of 7, (or the larger the value of a) the faster the time response of the actuator which in turns causes the higher the damping ratio of the Dutch roll mode. For a fairly large value of 7, (namely for a = 2), the damping ratio can only be obtained for a small value of gain ky. For a small value of 7., the high overshoot peak is due to the associated high gain.
Figure 2.46: Time response of y(t) and y(t) of the heading hold system of N250-100 aircraft
Figure 2.47: Effect of wind to the aircraft flight path The actual position of the aircraft has an offset angle A with respect to the main radial line and a radial distance of R with respect of VOR station. The aircraft position has a bearing y with respect to magnetic north and a distance d with respect to the VOR main bearing. From the geometry of the VOR guidance path the following relations can be readily obtained:
Figure 2.48: VOR guidance path geometry
Figure 2.49: The functional diagram of VOR-hold guidance-control system
Figure 2.50: The mathematical diagram of the VOR-hold guidance-control sys- tem
Figure 2.51: Navigation and guidance system: VOR hold The root locus of the above characteristic polynomial is illustrated in Fig.2.5: for the gain value of ky, < 0. From the root locus, if the following working point associated with ky = —1.1152 is taken, the following characteristic polynomia is obtained:
Figure 2.52: Root locus diagram for the VOR offset A(s) with respect to the bearing V,.¢ of the N250-100 aircraft
Figure 3.1: The functional diagram of the pitch damper system
Figure 3.3: Root locus of the inner control loop: pitch damper system of N250- 100 PA-2 at cruise condition with V = 250 KIAS and h = 15000’
Figure 3.4: Root locus of the inner control loop: pitch damper system of N250- 100 PA-2 at cruise condition with V = 250 KIAS and h = 15000’— enlarged around the phugoid mode
Figure 3.5: Time response of 0(t) and q(t) with and without pitch damper of N250-100 aircraft
Figure 3.6: The mathematical diagram of the pitch attitude hold with the pitch damper as the inner loop the outer characteristic equation of the phugoid damper is given by:
Figure 3.7: Root locus of outer control loop: phugoid damper of N250-100 PA2 at cruise condition
Figure 3.8: Root locus of outer control loop: phugoid damper of N250-100 PA2 at cruise condition—enlarged around phugoid mode
Figure 3.9: Time response of 6(t) of the phugoid damper with Kj = 3.7979 and K* = —0.2432
The components of the yaw damper system transfer function are: elimination in the steady state is to make the yaw damper not in opposition to the pilot input during the the stationary turning maneuver. If the r(t) still exists, it will work against the pilot and will make the coordinated turn difficult to achieve. Therefore, this yaw damper is designed only to enhance the dutch-roll damping in the transient condition such as the period of entering the gust or turbulence or the landing phase with ever decreasing speed.
Figure 3.11: The mathematical diagram of the yaw damper system with r — 6 feedback
where, the static sensitivity coefficient is given by: The transfer functions of the N250-100 due to the rudder input 6, is given as follows: As a case study, the application of the yaw damper to the N250-100 PA-2 Krincing Wesi aircraft will be presented. The aircraft is equipped with the yaw damper system to increase the ride quality particularly during the flight in turbulence and low-speed landing phase.
as a first order system as: The polynomial characteristic of the yaw damper system can be written as:
and the characteristic polynomial is:
Figure 3.13: Root locus of the yaw damper system with r —> 6, feedbac N250-100 aircraft—enlarged around the dutch-roll mode
It is evident from the root locus that the damping of dutch roll mode is improved while the time responses for the roll subsidence and spiral become faster. For sufficiently large gain of ky,, the dutch roll mode can be overdamped to break into two dutch roll subsidence modes.
by the following relations: The results of the calculated is presented in the following figure where three cases are compared. Note from the side slip angle response that, the application of the yaw damper enable a very effective suppresion of the response. In the yaw response, it is evident that for the damper system without the wash-out, che steady state y(t) is sustained constant at w,,(t) = 1 rad. Note, however chat, in the no yaw damper case, the yaw angle 7(t) can freely increase. The nclusion of the wash-out circuit enable the yaw damper to work only during che transient period. After the f,, signal is phased-out, the yaw angle w(t) can drift and increase without being inhibited by the yaw damper system. It can be concluded that by installing the wash-out circuit in the yaw damper, che maneuvers executed by the pilot during the steady state turn will not be »pposed by the yaw damper system. Fig.3.16 shows the phase-portrait of the ‘oll rate with respect to the roll angle. The effectivity of the yaw damper system n dampening down the dutch roll mode and the wash-out circuit in transmitting he yaw rate transiently is evident.
Figure 4.1: Minimum of a cost function J(x) at a stationary point
Table 3.1: Feedback for improving longitudinal stability servo has a task to receive command from the SAS computer only. In a num- ber of large size aircrafts, the AP and SAS servos are usually separated. The mathematical diagram of the pitch damper system is illustrated in Fig.3.2. If the SAS servo can be assumed to have a fast time response, it can be modeled as a constant gain:
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Theoretical and Applied Aerodynamics
by Gonzalo Anzaldo Muñoz
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CONTOH PROPOSAL PKM-GT (PROGRAM KREATIVITAS MAHASISWA GAGASAN TERTULIS) (Didanai Dikti 2017)
by Meda Aji Saputro
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Satellite Technology Principles and Applications Second Edition
by priyanka K
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A Modern Course in Aeroelasticity
by Gonzalo Anzaldo Muñoz
The fundamental questions arising in mechanics are: Why?, How?, and How much? The aim of this series is to provide lucid accounts written by authoritative researchers giving vision and insight in answering these questions on the subject... more
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Megaprojects and Risk: An Anatomy of Ambition
by Bent Flyvbjerg

2003

Back cover text: Megaprojects and Risk provides the first detailed examination of the phenomenon of megaprojects. It is a fascinating account of how the promoters of multi-billion dollar megaprojects systematically and self-servingly... more
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Propiedades de Materiales en Aeronautica
by Mayra Anzola
Reconocer, identificar y conocer las propiedades de los materiales es muy importante en el campo de la ingeniería, ya que son ellos los que aportan las características de resistencia y estabilidad de lo que se va a diseñar. Si no se... more
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Continuous and Discrete Time Signals and Systems
by Andrian Dwiputro
Continuous and Discrete Time Signals and SystemsContinuous and Discrete Time Signals and SystemsContinuous and Discrete Time Signals and SystemsContinuous and Discrete Time Signals and SystemsContinuous and Discrete Time Signals and... more
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The Jet Engines. Rolls-Royce publication
by Shuhrat Aslamshoev
|. 20-1 Thrust distribution of a typical single-spool axial flow engine.
Rolls-Royce Trent 800 Detailed engineering design began in 1988 to meet the propulsion requirements of the Airbus A330 (Trent 700) and Boeing 777 (Trent 800). The Trent first ran in August 1990, and in January 1994 a Trent 800 demonstrated a world record thrust of 106,087 Ib.
but it was eleven years before his engine completed its first flight. The Whittle engine formed the basis of the modern gas turbine engine, and from it was developed the Rolls-Royce Welland, Derwent, Nene and Dart engines. The Derwent and Nene turbo-jet engines had world-wide military applications; the Dart turbo-propeller engine became world famous as the power plant for the Vickers Viscount aircraft. Although other aircraft may be fitted; with later engines termed twin-spool, triple-spool, by-pass, ducted fan, unducted fan and propfan, these are inevitable developments of Whittle's early engine.
1-4 Hero’s engine - probably the earliest form of jet reaction.
of doing this, as described in Part 2, but in all instances the resultant reaction or thrust exerted on the engine is proportional to the mass or weight of air expelled by the engine and to the velocity change imparted to it. In other words, the same thrust can be provided either by giving a large mass of air a little extra velocity or a small mass of air a large extra
11. The ram jet engine (fig. 1-6) is an athodyd, or ‘aero-thermodynamic-duct to give it its full name. It has no major rotating parts and consists of a duct with a divergent entry and a convergent or
15. The mechanical arrangement of the gas turbine engine is simple, for it consists of only two main rotating parts, a compressor (Part 3) and a turbine (Part 5), and one or a number of combustion chambers (Part 4). The mechanical arrangement of various gas turbine engines is shown in fig. 1 -9. This simplicity, however, does not apply to all aspects of the engine, for as described in subsequent Parts the thermo and aerodynamic problems are somewhat complex. They result from the high operating tem- peratures of the combustion chamber and turbine, the effects of varying flows across the compressor
Fig. 1-9-2 Mechanical arrangement of gas turbine engines.
Fig. 1-10 Comparative propulsive efficiencies. 16. At aircraft speeds below approximately 450 miles per hour, the pure jet engine is less efficient than a propeller-type engine, since its propulsive efficiency depends largely on its forward speed; the pure turbo-jet engine is, therefore, most suitable for high forward speeds. The propeller efficiency does, however, decrease rapidly above 350 miles per hour due to the disturbance of the airflow caused by the high blade-tip speeds of the propeller. These charac- 17. The advantages of the propeller/turbine combination have to some extent been offset by the introduction of the by-pass, ducted fan and propfan engines. These engines deal with larger comparative airflows and lower jet velocities than the pure jet engine, thus giving a propulsive efficiency (Part 21) which is comparable to that of the turbo-prop and exceeds that of the pure jet engine (fig. 1-10).
18. The turbo/ram jet engine (fig. 1-11) combines the turbo-jet engine (which is used for speeds up to Mach 3) with the ram jet engine, which has good performance at high Mach numbers. ventional turbo-jet with the afterburner lit; at other flight conditions up to Mach 3, the afterburner is inoperative. As the aircraft accelerates through Mach 3, the turbo-jet is shut down and the intake air is diverted from the compressor, by guide vanes, and ducted straight into the afterburning jet pipe, which becomes a ram jet combustion chamber. This engine is suitable for an aircraft requiring high speed and
19. The engine is surrounded by a duct that has a variable intake at the front and an afterburning jet pipe with a variable nozzle at the rear. During take- off and acceleration, the engine functions as a con-
Fig. 2-1 Acomparison between the working cycle of a turbo-jet engine and a piston engine. Working cycle and airflow to use high octane fuels, in contrast to the low octane fuels and the light fabricated combustion chambers used on the turbine engine.
. 2-2. The working cycle on a pressure- volume diagram.
7. During the working cycle of the turbine engine, the airflow or 'working fluid' receives and gives up heat, so producing changes in its pressure, volume and temperature. These changes as they occur are closely related, for they follow a common principle that is embodied in a combination of the laws of Boyle and Charles. Briefly, this means that the product of the pressure and the volume of the air at the various stages in the working cycle is proportion- al to the absolute temperature of the air at those 6. The use of air-cooled blades in the turbine assembly permits a higher gas temperature and a consequently higher thermal efficiency. y. 2-3. An airflow through divergent and convergent ducts.
14. The design of the passages and nozzles is of great importance, for upon their good design will depend the efficiency with which the energy changes are effected. Any interference with the smooth airflow creates a loss in efficiency and could result in component failure due to vibration caused by eddies or turbulence of the airflow.
TWIN-SPOOL AXIAL FLOW CONTRA-ROTATING REAR FAN (with free power turbine) Fig, 2-5-2 Airflow systems. Working cycle and airflow
TRIPLE-SPOOL AXIAL FLOW FRONT FAN TURBO JET ENGINE (high by-pass ratio)
INTRODUCTION 1. In the gas turbine engine, compression of the air before expansion through the turbine is effected by one of two basic types of compressor, one giving centrifugal flow and the other axial flow. Both types are driven by the engine turbine and are usually coupled direct to the turbine shaft. 2. The centrifugal flow compressor (fig. 3-1) is a single or two stage unit employing an impeller to
. 3-1. Atypical centrifugal flow compressor.
8. To maximize the airflow and pressure rise through the compressor requires the impeller to be rotated at high speed, therefore impellers are designed to operate at tip speeds of up to 1,600 ft.
Construction 9. To maintain the efficiency of the compressor, it is necessary to prevent excessive air leakage between the impeller and the casing; this is achieved by keeping their clearances as small as possible (fig. 3- 4).
vanes are in line with the direction of the resultant airflow from the impeller (fig. 3-6). The clearance between the impeller and the diffuser is an important factor, as too small a clearance will set up aerodynamic buffeting impulses that could be transferred to the impeller and create an unsteady airflow and vibration. Fig. 3-6 Airflow at entry to diffuser.
ig. 3-5 Typical impellers for centrifugal compressors.
the rotor shaft and the stator casing. This is necessary to maintain a near constant air axial velocity as the density increases through the length of the compressor. The convergence of the air annulus is achieved by the tapering of the casing or rotor. A combination of both is also possible, with the arrangement being influenced by manufacturing problems and other mechanical design factors.
19. During operation the rotor is turned at high speed by the turbine so that air is continuously induced into the compressor, which is then accelerated by the rotating blades and swept rearwards onto the adjacent row of stator vanes. The pressure rise results from the energy imparted to the air in the rotor which increases the air velocity. The air is then decelerated (diffused) in the following
Fig. 3-13 Methods of securing vanes to compressor casing. 29. If the operating conditions imposed upon the compressor blade departs too far from the design intention, breakdown of airflow and/or aerodynami- cally induced vibration will occur. These phenomena may take one of two forms; the blades may stall because the angle of incidence of the air relative to the blade is too high (positive incidence stall) or too low (negative incidence stall). The former is a front stage problem at low speeds and the latter usually affects the rear stages at high speed, either can lead to blade vibration which can induce rapid destruction. If the engine demands a pressure rise from the compressor, which is higher than the blading can sustain, ‘surge’ occurs. In this case there is an instan- taneous breakdown of flow through the machine and the high pressure air in the combustion system is expelled forward through the compressor with a loud ‘pang’ and a resultant loss of engine thrust.
30. Where high pressure ratios on a single shaft are required it becomes necessary to introduce airflow control into the compressor design. This may take 31. Materials are chosen to achieve the most cost effective design for the components in question, in practice for aero engine design this need is usually best satisfied by the lightest design that technology allows for the given loads and temperatures prevailing.
Fig. 3-17 A pneumatically operated bleed valve system.
Fig. 3-18 An electronically operated bleed valve system.
Fig. 4-1 An early combustion chamber.
9. The temperature of the gases released by combustion is about 1,800 to 2,000 deg. C., which is far too hot for entry to the nozzle guide vanes of the turbine. The air not used for combustion, which amounts to about 60 per cent of the total airflow, is therefore introduced progressively into the flame tube. Approximately a third of this is used to lower the gas temperature in the dilution zone before it enters
injected fuel droplets by rapidly bringing them to ignition temperature. the turbine and the remainder is used for cooling the walls of the flame tube. This is achieved by a film of cooling air flowing along the inside surface of the flame tube wall, insulating it from the hot combustion gases (fig. 4-4). A recent development allows cooling air to enter a network of passages within the flame tube wall before exiting to form an insulating film of air, this can reduce the required wall cooling airflow by up to 50 per cent. Combustion should be completed before the dilution air enters the flame tube, otherwise the incoming air will cool the flame and incomplete combustion will result.
4-4 Flame tube cooling methods.
chamber has an inner flame tube around which there is an air casing. The air passes through the flame tube snout and also between the tube and the outer casing as already described in para. 6. 17. The separate flame tubes are all interconnect- ed. This allows each tube to operate at the same pressure and also allows combustion to propagate around the flame tubes during engine starting.
20. The main advantage of the annular chamber is that, for the same power output, the length of the chamber is only 75 per cent of that of a tubo-annular system of the same diameter, resulting in consider- able saving of weight and production cost. Another advantage is the elimination of combustion propagation problems from chamber to chamber. 18. The tubo-annular combustion chamber bridges the evolutionary gap between the multiple and annular types. A number of flame tubes are fitted inside a common air casing (fig. 4-8). The airflow is similar to that already described. This arrangement combines the ease of overhaul and testing of the multiple system with the compactness of the annular system.
4-8 Tubo-annular combustion chamber.
and compact gas turbine combustion chamber must release heat at exceptionally high rates. another way this is an expenditure of potential heat at a rate equivalent to approximately 150,000 horse- power. 26. For example, at take-off conditions a Rolls- Royce RB211-524 engine will consume 20,635 Ib. of fuel per hour. The fuel has a calorific value of approx- imately 18,550 British thermal units per Ib., therefore the combustion chamber releases nearly 106,300 British thermal units per second. Expressed in
Fig. 4-11 Combustion stability limits. 34. Oxides of nitrogen are formed under the same conditions as those required for the suppression of the other pollutants, Therefore it is desirable to cool the flame as quickly as possible and to reduce the time available for combustion. This conflict of conditions requires a compromise to be made, but continuing improvements in combustor design and performance has led to a substantially ‘cleaner’ combustion process.
5-1 Atriple-stage turbine with single shaft system. from the gas flow, the rotational speed at which it must be produced and the diameter of turbine permitted. ratio fan engines that feature an intermediate pressure system, another turbine may be interposed between the high and low pressure turbines, thus forming a triple-spool system (fig, 5-3). On some engines, driving torque is derived from a free-power turbine (fig. 5-4). This method allows the turbine to run at its optimum speed because it is mechanically independent of other turbine and compressor shafts.
Fig. 5-2. A twin turbine and shaft arrangement. 5. The design of the nozzle guide vane and turbine blade passages is based broadly on aerodynamic considerations, and to obtain optimum efficiency, compatible with compressor and combustion design, the nozzle guide vanes and turbine blades are of a 4. The mean blade speed of a turbine has consid- erable effect on the maximum efficiency possible for a given stage output. For a given output the gas velocities, deflections, and hence losses, are reduced in proportion to the square of higher mean blade speeds. Stress in the turbine disc increases as the square of the speed, therefore to maintain the same stress level at higher speed the sectional thickness, hence the weight, must be increased dis- proportionately. For this reason, the final design is a compromise between efficiency and weight. Engines
reaction turbine blades but the impulse-reaction combination (fig. 5-5). The proportion of each principle incorporated in the design of a turbine is largely dependent on the type of engine in which the turbine is to operate, but in general it is about 50 per cent impulse and 50 per cent reaction. Impulse-type turbines are used for cartridge and air starters (Part 11). basic aerofoil shape. There are three types of turbine; impulse, reaction and a combination of the two known as impulse-reaction. In the impulse type the total pressure drop across each stage occurs in the fixed nozzle guide vanes which, because of their convergent shape, increase the gas velocity whilst reducing the pressure. The gas is directed onto the turbine blades which experience an impulse force caused by the impact of the gas on the blades. In the reaction type the fixed nozzle guide vanes are designed to alter the gas flow direction without changing the pressure. The converging blade passages experience a reaction force resulting from the expansion and acceleration of the gas. Normally gas turbine engines do not use pure impulse or pure
|. 5-5 Comparison between a pure Impulse turbine and an impulse/reaction turbine.
13. The nozzle guide vanes are of an aerofoil shape with the passage between adjacent vanes forming a convergent duct. The vanes are located (fig. 5-8) in the turbine casing in a manner that allows for expansion. g. 5-7 Gas flow pattern through nozzle and blade.
-ig. 5-8 Typical nozzle guide vanes showing their shape and location.
Fig. 5-9 Various methods of attaching blades to turbine discs. 24. Among the obstacles in the way of using higher turbine entry temperatures have always been the effects of these temperatures on the nozzle guide vanes and turbine blades, The high speed of rotation which imparts tensile stress to the turbine disc and blades is also a limiting factor. 21. Fig. 5-10 shows a twelve stage contra-rotating free power turbine driving a contra-rotating rear fan. This design has only one row of static nozzle guide vanes. The remaining nozzle guide vanes are, in effect, turbine blades attached to a rotating casing which revolves in the opposite direction to a rotating drum. Since all but one aerofoil row extracts energy from the gas stream, contra-rotating turbines are
Fig. 5-10 Free power contra-rotating turbine.
Nozzle guide vanes
Fig. 5-13 Comparison of turbine blade life properties.
|. 6-2 Exhaust system with thrust reverser, noise suppressor and two position propelling nozzle.
3 Gas flow through a convergent- divergent nozzle. 5. The exhaust gases pass to atmosphere through the propelling nozzle, which is a convergent duct, thus increasing the gas velocity (Part 2). In a turbo- jet engine, the exit velocity of the exhaust gases is subsonic at low thrust conditions only. During most operating conditions, the exit velocity reaches the speed of sound in relation to the exhaust gas
Fig. 6-5 High by-pass ratio engine exhaust systems. with the exit proper now being at the end of the flared divergent section. When the gas enters the convergent section of the nozzle, the gas velocity increases with a corresponding fall in static pressure. The gas velocity at the throat corresponds to the local sonic velocity. As the gas leaves the restriction of the throat and flows into the divergent section, it progressively increases in velocity towards the exit. The reaction to this further increase in momentum is a pressure force acting on the inner wall of the nozzle. A component of this force acting parallel to the longitudinal axis of the nozzle produces the further increase in thrust. 7. From the illustration (fig. 6-3), it will be seen that the convergent section exit now becomes the throat,
The Ghost was designed as a larger and more powerful version of the Goblin. After running for the first time on 2 September 1945 the engine was cleared for flight in the outer nacelles of an Avro Lancastrian at 4000 Ib thrust. The Ghost later went into production at 5000 Ib thrust to power the de Havilland Comet 1 airliner and Venom fighter.
the compressor and the conditions at the outlet from the compressor. Since the pressure and the velocity at the inlet to the compressor are zero, it is only necessary to consider the force at the outlet from the compressor. Therefore, given that the compressor-
accessory unit to be located in any desirable location around the engine and decreased the power transmitted through individual gears, it necessitated a large internal gearbox. Additionally, numerous radial driveshafts had to be incorporated within the design. This led to an excessive amount of time required for disassembly and assembly of the engine for maintenance purposes. rotate to provide the same power. However, this raises the internal stress and gives greater dynamic problems which result in vibration. A long radial driveshaft usually requires a roller bearing situated halfway along its length to give smooth running. This allows a rotational speed of approximately 25,000 r.p.m. to be achieved with a shaft diameter of less than 1.5 inch without encountering serious vibration problems.
y. 7-4 An external gearbox and accessory units.
INTRODUCTION 1. The lubrication system is required to provide lubrication and cooling for all gears, bearings and splines. It must also be capable of collecting foreign matter which, if left in a bearing housing or gearbox, can cause rapid failure. Additionally, the oil must protect the lubricated components which are manu- factured from non-corrosion resistant materials. The oil must accomplish these tasks without significant deterioration.
. 8-1 Apressure relief valve type oil system. 6 Fig. 8-1 shows the pressure relief valve system for a turbo-propeller engine and indicates the basic components that comprise an engine lubrication system. The oil pressure pump draws oil from the tank through a strainer which protects the pump gears from debris which may have entered the tank, Oil is then delivered through a pressure filter to the pressure relief valve which maintains a constant oil delivery pressure to the feed jets in the bearing chambers. Some engines may have an additional relief valve (pressure limiting valve) which is fitted at the oil pressure pump outlet. This valve is set to open at a much higher value than the pressure relief valve to return the oil to the inlet side of the oil pressure pump in the event of the system becoming blocked. Asimilar valve may also be fitted across the pressure filter to prevent oil starvation of the bearing chambers should the filter become partially blocked or the oil having a high viscosity under cold starting conditions preventing sufficient flow through the filter. Provision is also made to supply oil to the propeller pitch control system, reduction gear and torquemeter 5. In the pressure relief valve system the oil flow to the bearing chambers is controlled by limiting the pressure in the feed line to a given design value. This is accomplished by the use of a spring loaded valve which allows oil to be directly returned from the pressure pump outlet to the oil tank, or pressure pump inlet, when the design value is exceeded. The valve opens at a pressure which corresponds to the idling speed of the engine, thus giving a constant feed pressure over normal engine operating speeds. However, increasing engine speed causes the bearing chamber pressure to rise sharply. This reduces the pressure difference between the bearing chamber and feed jet, thus decreasing the oil flow rate to the bearings as engine speed increases. To alleviate this problem, some pressure relief valve systems use the increasing bearing chamber pressure to augment the relief valve spring load, This maintains a constant flow rate at the higher engine speeds by increasing the pressure in the feed line as the bearing chamber pressure increases.
8. The full flow system achieves the desired oil flow rates throughout the complete engine speed range by dispensing with the pressure relief valve and allowing the pressure pump delivery pressure to supply directly the oil feed jets. Fig. 8-2 shows an example of this system which may be found on a turbo-fan engine. The pressure pump size is determined by the flow required at maximum engine speed. The use of this system allows smaller pressure and scavenge pumps to be used since the large volume of oil which is spilled by the pressure relief valve system at maximum engine speed is obviated. 7. Although the pressure relief valve system operates satisfactorily for engines which have a low bearing chamber pressure, which does not unduly increase with engine speed, it becomes an undesirable system for engines which have high chamber pressures. For example, if a bearing chamber has a maximum pressure of 90 Ib. per sq. in. It would require a pressure relief valve setting of 130 Ib. per sq. in. to produce a pressure drop of 40 lb. per sq. in. at the oil feed jet. This results in the
8-3 A total loss (expendable) oil system. the total loss oil system is generally used. The system is simple and incurs low weight penalties because it requires no oil cooler, scavenge pump or filters. On some engines oil is delivered in a continuous flow to the bearings by a plunger-type pump, indirectly driven from the compressor shaft; on others it is delivered by a piston-type pump operated by fuel pressure (fig. 8-3). In the latter, the oil supply is automatically selected by the high pressure fuel shut-off valve (cock) during engine starting and is delivered as a single shot to the front and rear bearings. On some engines provision is made for a 10. For engines which run for periods of short duration, such as booster and vertical lift engines,
12. The oil tank (fig. 8-4) is usually mounted on the engine and is normally a separate unit although it may also be an integral part of the external gearbox. It must have provision to allow the lubrication system to be drained and replenished. A sight glass or a dipstick must also be incorporated to allow the oil 11. After lubricating the fuel unit and front bearings, the oil from the front bearing drains into a collector tray and is then ejected into the main gas stream through an ejector nozzle. The oil that has passed through the rear bearing, drains into a reservoir at the rear of the bearing where it is retained by centrifugal force until the engine is shut down. This
Fig. 8-6 Asingle-shot oil pump.
Fig. 8-11 Athread-type oil filter.
debris from each bearing chamber. They are basically permanent magnets inserted in the oil flow and are retained in self-sealing valve housings. Safety features incorporated in the design ensure correct retention within the housing. Upon examination they can provide a warning of impending failure without having to remove and inspect the filters. They are designed to be removed during maintenance inspection, for condition, monitoring purposes (Part 24), without oil loss occurring. Additionally they may be connected to a cockpit warning system to give an in-flight indication.
Fig. 8-12 Atypical pressure and scavenge filter.
14. Both the thermocouple and thermistor detectors have properties making them ideally suited to this application. The thermocouple comprises two dissimilar metals which are joined together to form two junctions. As the temperature difference between the two junctions increases an E.M.F. is produced in the circuit and it is this E.M.F. that triggers the fire 10. Fig. 14-2 shows a more complex cooling and ventilation system used on a turbo-fan engine. Air is induced from the intake duct and also delivered from the fan to provide multi-zone cooling, each zone having its own calibrated cooling flow.
5. High thermal efficiency is dependent upon high turbine entry temperature, which is limited by the turbine blade and nozzle guide vane materials. Continuous cooling of these components allows their environmental operating temperature to exceed the material's melting point without affecting the blade and vane integrity. Heat conduction from the turbine blades to the turbine disc requires the discs to be cooled and thus prevent thermal fatigue and uncon- trolled expansion and contraction rates. 3. Animportant consideration at the design stage of a gas turbine engine is the need to ensure that certain parts of the engine, and in some instances certain accessories, do not absorb heat to the extent that is detrimental to their safe operation. The principal areas which require air cooling are the combustor and turbine. Refer to Part 4 for combustor cooling techniques. that is detrimental to their safe operation. The
g. 9-3 Development of high pressure turbine blade cooling. 7. The 'pre-swirl nozzles' (fig. 9-2) reduce the temperature and pressure of the cooling air fed to the disc for blade cooling. The nozzles also impart a 9. Air cooling of the engine bearing chambers is not normally necessary since the lubrication system (Part 8) is adequate for cooling purposes. Additionally, bearing chambers are located, where possible, in the cooler regions of the engine. In instances where additional cooling is required, it is good practice to have a double skinned bearing housing with cooling air fed into the intermediate space.
‘ig. 9-4 High pressure nozzle guide vane construction and cooling.
g.9-5 Ahypothetical turbine cooling and sealing arrangement.
Fig. 10-3 Simplified fuel systems for turbo-propeller and turbo-jet engines.
Fig. 10-4 A pressure control system (turbo-propeller engine).
-ig. 10-5 A pressure control system (turbo-jet engine).
Fig. 10-6 Acceleration control by dashpot throttle.
Fig. 10-8 Servo pressure control by kinetic valve.
Fig. 10-10 Governor movement and fuel flow control.
Fig. 10-11 Effect of water reset on speed control governor. 66. The F.F.R. is driven by the engine through a gear train and has two rotating valves. One valve, 69. When the throttle is slowly opened, the throttle- controlled orifice is increased and the control pressure falls, thus allowing the pressure ratio control valve to move towards the closed position (acceleration stop). F.F.R. capsule chamber pressure
Fig. 10-15 Various stages of fuel atomization.
Fig. 10-16 A Simplex fuel spray nozzle.
96. The spill spray nozzle system, however, involves a somewhat modified type of fuel supply and control system from that used with the previous types. A means has to be provided for removing the
Fig. 10-18 A Duple fuel spray nozzle and pressurizing valve.
106. Fuels for aircraft gas turbine engines must conform to strict requirements to give optimum engine performance, economy, safety and overhaul life. Fuels are classed under two headings, kerosine- type fuel and wide-cut gasoline-type fuel.
Fig. 10-21 Relationship between calorific value and specific gravity. Other factors which affect the choice of heat per unit of volume or weight, must also be taken into consid- eration; these include the type of aircraft, the duration of flight, and the required balance between fuel weight and payload.
2. The functioning of both systems is co-ordinatec during a starting cycle and their operation is auto- matically controlled after the initiation of the cycle by an electrical circuit. A typical sequence of events during the start of a turbo-jet engine is shown in fig. 11-1.
6. The electrical supply may be of a high or low voltage and is passed through a system of relays and resistances to allow the full voltage to be progres- sively built up as the starter gains speed. It also provides the power for the operation of the ignition system. The electrical supply is automatically cancelled when the starter load is reduced after the engine has satisfactorily started or when the time cycle is completed. A typical electrical starting system is shown in fig. 11-3. 4. The starter motor must produce a high torque and transmit it to the engine rotating assembly in a manner that provides smooth acceleration from rest up to a speed at which the gas flow through the
-ig. 11-3 A low voltage electrical starting system.
Gas turbine rotation of the starter ceases. A typical air starting system is shown in fig. 11-7.
operates and shuts down the gas turbine starter. As the starter runs down, the clutch automatically disengages from the output shaft and the main engine accelerates up to idling r.p.m. under its own power. cycle the pump /starter functions as a normal hydraulic pump.
23. Operation of the transistorized ignition unit is similar to that of the D.C. trembler-operated unit, except that the trembler-unit is replaced by a transistor chopper circuit. A typical transistorized unit is shown in fig. 11-11; such a unit has many advantages over the trembler-operated unit because it has no moving parts and gives a much longer operating life. The size of the transistorized unit is reduced and its weight is less than that of the trembler-operated unit. 22. The ignition unit shown in fig. 11-10 is atypical D.C. trembler-operated unit. An induction coil, operated by the trembler mechanism, charges the reservoir capacitor (condenser) through a high voltage rectifier. When the voltage in the capacitor is equal to the breakdown value of a sealed discharge gap, the energy is discharged across the face of the igniter plug. A choke is fitted to extend the duration of the discharge and a discharge resistor is fitted to
24. The A.C. ignition unit, shown in fig, 11-12, receives an alternating current which is passed through a transformer and rectifier to charge a capacitor. When the voltage in the capacitor is equal to the breakdown value of a sealed discharge gap, the capacitor discharges the energy across the face of the igniter plug. Safety and discharge resistors are fitted as in the trembler-operated unit.
Fig. 12-2 Diagrammatic arrangement of engine control and instrumentation. 6. On a turbo-propeller engine, the throttle lever is interconnected with the propeller control unit (P.C.U.), thus maintaining single lever operation of the engine. Similarly, the throttle control lever of a helicopter is interconnected with the collective pitch lever, so ensuring that an increase in pitch is accompanied by an increase in engine power, 5. On engines fitted with afterburning, single lever control is maintained, although a further fuel system is required to supply and control the fuel to the afterburner (Part 16). Controls and instrumentation
11. In each instance, an indication of thrust output is given, although when only the turbine discharge pressure is measured, correction is necessary for variation of inlet pressure; however, both types may require correction for variation of ambient air temperature. To compensate for ambient atmospheric conditions, it is possible to set a correction figure to a sub-scale on the gauge; thus, 14. E.P.R. can be indicated by either electro- mechanical or electronic transmitters. In both cases the inputs to the transmitter are engine inlet pressure (P1) and an integrated pressure (PINT) comprised of
Fig. 12-6 Variable-reluctance speed probe and phonic wheel.
‘ig. 12-8 Atypical double element thermocouple system. 43. Although the amount of fuel consumed during a given flight may vary slightly between engines of the same type, fuel flow does provide a useful indication of the satisfactory operation of the engine and of the amount of fuel being consumed during the flight. A typical system consists of a fuel flow transmitter, which is fitted into the low pressure fuel system, and an indicator, which shows the rate of fuel flow and the 41. The temperature and pressure of the low pressure fuel supply are electrically transmitted to their respective indicators and these show if the low pressure system is providing an adequate supply of fuel without cavitation and at a temperature to suit the operating conditions. The fuel temperature and
Fig. 12-12 Typical electronic indicating display. 55. Only those parameters required by the crew to set and monitor engine thrust are permanently displayed on the screen. The system monitors the remaining parameters and displays them only if one or more exceed safe limitations. The pilot can, however, override the system and elect to have all main parameters in view at any time (fig. 12-12).
multitude of instruments traditionally present are replaced by the C.R.T.'s. Controls and instrumentation
7. Ifthe nose cone is anti-iced its hot air supply may be independent or integral with that of the nose cowl and compressor stators. For an independent system, the nose cone is usually anti-iced by a continuous 5. The hot air system provides surface heating of the engine and/or powerplant where ice is likely to form. The protection of rotor blades is rarely necessary, because any ice accretions are dispersed by centrifugal action. If stators are fitted upstream of the first rotating compressor stage these may require ig. 13-3 Combination of hot air, oil and electrical ice protection.
forming, but the intermittently heated areas allow ice to form, during their 'heat-off period. During the 'heat- on' period, adhesion of the ice is broken and it is then removed by aerodynamic forces. without causing any run-back icing to occur behind the heated areas.
1. All gas turbine engines and their associated installation systems incorporate features that minimize the possibility of an engine fire. It is essential, however, that if a failure does take place and results in a fire, there is provision for the immediate detection and rapid extinction of the fire, and for the prevention of it spreading. The detection and extinguishing systems must add as little weight to the installation as possible.
‘ig. 14-2 Cooling and ventilation - turbo-fan engine.
Therefore given that the turbine-- This negative value means a force acting in a rearward direction.
16. The gas filled detector consists of stainless steel tubing filled with gas absorbent material and in the event of a fire or overheat condition the temperature rise will cause the core of the sensing loop to expel the absorbed active gas into the sealed tube causing a rapid increase in pressure. This build up of pressure is sensed by the detector alarm switch. Should the sensing loop become damaged causing a loss of the pressurized gas, an integrity switch will indicate a detection loop fault on the appropriate engine. Fire indication is given by a warning light and bell. 14-3 A continuous element fire detecting system. 19. An engine fire must be contained within the power plant and not be allowed to spread to other parts of the aircraft. The cowlings that surround the engine are usually made of aluminium alloys, which would be unable to contain a fire when the aircraft is static. During flight, however, the airflow around the cowlings provides sufficient cooling to render them fireproof. Fireproof bulkheads and any cowlings that are not affected by a cooling airflow, and sections of cowlings around certain outlets that may act as ‘flame-holders', are usually manufactured from steel or titanium.
14-4 Atypical fire extinguishing system.
Fig. 15-1 Comparative landing runs with and without thrust reversal.
Fig. 15-5 A propeller pitch control system.
Fig. 15-7 Acold stream thrust reverser installation. 20. The clamshell and bucket target doors (fig. 15- 6) described in paras. 9 and 12 form part of the jet 22. Both the clamshell door system and the bucket target system are subjected to high temperatures and to high gas loads. The components of both systems, especially the doors, are therefore
16: Afterburning
Fig. 16-2 Examples of afterburning jet pipes and propelling nozzles.
8. An atomized fuel spray is fed into the jet pipe through a number of burners, which are so arranged as to distribute the fuel evenly over the flame area. Combustion is then initiated by a catalytic igniter, which creates a flame as a result of the chemical reaction of the fuel/air mixture being sprayed on to a platinum-based element, by an igniter plug adjacent to the burner, or by a hot streak of flame that originates in the engine combustion chamber (fig. 9. In view of the high temperature of the gases entering the jet pipe from the turbine, it might be assumed that the mixture would ignite spontaneous- ly. This is not so, for although cool flames form at
. 16-4 Typical afterburning jet pipe equipment.
hydraulic medium, but some systems use fuel. Nozzle movement is achieved by the hydraulic operating rams which are pressurized by an oil pump, pump output being controlled by a linkage from the pressure ratio control unit. When an increase in afterburning is selected, the afterburner fuel control unit schedules an increase in fuel pump output. The jet pipe pressure (P6) increases, altering the pressure ratio across the turbine (P3/P6). The pressure ratio control unit alters oil pump output, causing an out-of-balance condition between the hydraulic ram load and the gas load on the nozzle flaps. The gas load opens the nozzle to increase its exit area and, as the nozzle opens, the increase in nozzle area restores the P3/P6 ratio and the pressure ratio control unit alters oil pump output until balance is restored between the hydraulic rams and the gas loading on the nozzle flaps.
‘ig. 16-6 Asimplified typical afterburner fuel control system.
16-8 Thrust increase and temperature ratio.
‘ig. 16-10 Afterburning and its effect on the rate of climb. 23. Afterburning always incurs an increase in specific fuel consumption and is, therefore, generally limited to periods of short duration. Additional fuel must be added to the gas stream to obtain the required temperature ratio (para. 19). Since the temperature rise does not occur at the peak of compression, the fuel is not burnt as efficiently as in the engine combustion chamber and a_ higher specific fuel consumption must result. For example, assuming a specific fuel consumption without after- burning of 1,15 Ib./hr./lb. thrust at sea level and a speed of Mach 0,9 as shown in fig. 16-9. then with 70 per cent afterburning under the same conditions of flight, the consumption will be increased to
17-3 Atypical compressor inlet injection system.
. 17-4 Atypical combustion chamber injection system.
Rolls-Royce Pegasus
Fig. 18-1 Michel Wibault's ground attack gyropter (concept) 1956
each side of the aircraft. A further development was to use the fan to supercharge the engine, exhausting the by-pass air through one pair of swivelling nozzles and adding a second pair of swivelling nozzles to the 4. The principle proposed by M. Wibault was developed by using a pure jet engine with a free power turbine to drive an axial flow fan which exhausted into a pair of swivelling nozzles, one on
8. Thrust deflection on a single nozzle is accom- plished by connecting together sections of the jet
9. The two and four nozzle deflector systems use side mounted nozzles (fig. 18-6) which can rotate on simple bearings through an angle of well over 90 degrees so that reverse thrust can be provided if required. A simple drive system, for example, a sprocket and chain, can be used and by mechanical connections all the nozzles can be made to deflect simultaneously. For forward flight, to avoid a high performance loss and consequent increase in fuel consumption, careful design of the exhaust unit and nozzle aerodynamic passages are essential to minimize the pressure losses due to turning the exhaust flow through two close coupled bends (fig. 18-7).
15. Lift engines can be designed to operate in the vertical or horizontal position and a thrust deflecting nozzle fitted to provide some of the advantages of thrust vectoring. Alternatively, the engine may be mounted so that it can swivel through a large angle to provide thrust vectoring. The lift-jet engine will have an extremely hot, high velocity jet exhaust and to reduce ground erosion by the jet the normal 16. The lift-fan engine is designed to reduce the jet exhaust velocity, to reduce ground erosion and allow operation from unprepared ground surfaces. It also reduces the jet noise significantly. A range of design options have been considered for this type of engine and some are shown on fig. 18-11.
exhaust nozzle may be replaced by a multi-lobe nozzle to increase the rate of mixing with the surrounding air.
Fig. 18-11 _Lift-fan engine configurations. 17. Direct lift remote systems duct the by-pass air or engine exhaust air to downward facing lift nozzles remote from the engine. These nozzles may be in the front fuselage of the aircraft or in the wings. The engine duct is blocked by means of a diverter similar to that described in para. 10. 19. The advantage of the remote lift system is that it gives some freedom to the aircraft to position the
18. The remote lift-fan (fig. 18-12) is mounted in the aircraft wing or fuselage, and is driven mechanically or by air or gas ducted into a tip turbine, The drive system is provided by the main propulsion power plant or by a separate engine. Vertical/short take-off and landing
20. This method consists of having propulsion engines which can be mechanically swiveled closed LIFT THRUST AUGMENTATION
propulsion system to the best advantage whilst still maintaining the resultant thrust near the aircraft centre of gravity in the jet lift mode. This freedom is achieved at a cost of increased volume, particularly with the gas driven systems, due to the size of the ducts to feed the gas to the remote lift system. Although the mechanically driven remote lift-fan eliminates the need for these large gas ducts, it is done at the expense of long shafts and high power gearboxes and clutch systems.
Fig. 18-18 Reaction control system. AIRCRAFT CONTROL
PLAIN NOZZLE (low mixing rate) HIGH NOISE LEVEL “ig. 19-3 Change of exhaust jet pattern to reduce noise level.
5. Jet exhaust noise is caused by the violent and hence extremely turbulent mixing of the exhaust gases with the atmosphere and is influenced by the shearing action caused by the relative speed between the exhaust jet and the atmosphere. The small eddies created near the exhaust duct cause high frequency noise but downstream of the exhaust jet the larger eddies create low frequency noise. Additionally, when the exhaust jet velocity exceeds the local speed of sound, a regular shock pattern is formed within the exhaust jet core. This produces a discrete (single frequency) tone and selective ampli- fication of the mixing noise, as shown in fig. 19-2. A reduction in noise level occurs if the mixing rate is accelerated or if the velocity of the exhaust jet relative to the atmosphere is reduced. This can be achieved by changing the pattern of the exhaust jet as shown in fig. 19-3. 6. Compressor and turbine noise results from the interaction of pressure fields and turbulent wakes from rotating blades and stationary vanes, and can be defined as two distinct types of noise; discrete tone (single frequency) and broadband (a wide range of frequencies). Discrete tones are produced by the regular passage of blade wakes over the stages downstream causing a series of tones and harmonics from each stage. The wake intensity is largely dependent upon the distance between the rows of blades and vanes. If the distance is short then there is an intense pressure field interaction which results in a strong tone being generated. With the high bypass engine, the low pressure compressor (fan) blade wakes passing over downstream vanes produce such tones, but of a lower intensity due to lower velocities and larger blade/vane separations. Broadband noise is
‘ig. 19-4 Comparative noise sources of low and high by-pass engines. 8. The introduction of a single stage low pressure compressor (fan) significantly reduces the compressor noise because the overall turbulence and interaction levels are diminished. When the by- pass ratio is in excess of approximately 5 to 1, the jet exhaust noise has reduced to such a level that the increased internal noise source is predominant. A comparison between low and high by-pass engine noise sources is shown in fig. 19-4. 10. Noise suppression of internal sources is approached in two ways; by basic design to minimize noise originating within or propagating from the engine, and by the use of acoustically absorbent linings. Noise can be minimized by reducing airflow disruption which causes turbulence. This is achieved by using minimal rotational and airflow velocities and reducing the wake intensity by appropriate spacing between the blades and vanes. The ratio between the number of rotating blades and stationary vanes can also be advantageously employed to contain noise within the engine.
Fig. 19-5 Types of noise suppressor.
a good understanding of the mechanisms of noise generation and comprehensive noise design rules exist. As previously indicated, these are founded on the need to minimize turbulence levels in the airflow, reduce the strength of interactions between rotating blades and stationary vanes, and the optimum use of acoustically absorbent linings.
‘ig. 20-2 A hypothetical combustion chamber showing values required for calculating thrust. = 11,158 Ib., the same as previously calculated 23. When the engine is fitted with an afterburner (Part 16), the gases passing through the exhaust
Therefore, assuming values of mass flow, pressure and area to be the same as in the previous calcula- tions i.e.
system are reheated to provide additional thrust. The effect of afterburning is to increase the volume of the exhaust gases, thus producing a higher exit velocity at the propelling nozzle.
24. Assuming that an afterburner jet pipe and propelling nozzle are fitted to the engine used in the previous calculations, and the new conditions at the propelling nozzle are as follows- Therefore, compared with the previous calculation in para. 17, it will be seen that the negative thrust is reduced from -5,587 Ib. to -2,676 Ib.; the overall positive thrust is thus increased by 2,911 |b; which is equivalent to a thrust increase of more than 25 per cent.
Fig. 21-1 Temperature and pressure notation of a typical turbo-jet engine. 7. To enable the performance of similar engines to be compared, it is necessary to standardize in some conventional form the variations of air temperature and pressure that occur with altitude and climatic conditions. There are in use several different definitions of standard atmospheres, the one in most common use being the International Standard Atmosphere (I.S.A.). This is based on a temperature lapse rate of approximately 1.98 K. degrees per 1,000ft,, resulting in a fall from 288.15 deg.K. (15 deg.C) at sea level to 216.65 deg.K (-56.5 deg.C.) at 36,089 ft. (the tropopause). Above this altitude the 6. These and other factors are presented in curves and graphs, calculated from the basic gas laws (Part 2), and are proved in practice by bench and flight testing, or by simulating flight conditions in a high altitude test cell. To make these calculations, specific symbols are used to denote the pressures and tem- peratures at various locations through the engine; for
Thus at 600 m.p.h. one lb. of thrust is the equivalent of about 3 t.h.p. 18. From the definitions and formulae stated in para, 17; under flight conditions, the net thrust of the
20. From the thrust equation in para. 18, it is apparent that if the jet velocity remains constant, independent of aircraft speed, then as the aircraft speed increases the thrust would decrease in direct proportion. However, due to the 'ram ratio' effect from the aircraft forward speed, extra air is taken into the engine so that the mass airflow and also the jet velocity increase with aircraft speed. The effect of this tends to offset the extra intake momentum drag
ig. 21-5 The effect of aircraft speed on s.h.p. and fuel consumption.
Fig. 21-6 The effects of altitude on thrust and fuel consumption.
38. The efficiency of conversion of fuel energy to kinetic energy is termed thermal or internal efficiency and, like all heat engines, is controlled by the cycle pressure ratio and combustion temperature. Unfortunately, this temperature is limited by the thermal and mechanical stresses that can be tolerated by the turbine. The development of new materials and techniques to minimize these limitations is continually being pursued. Fig. 21-8 The effect of air temperature on a typical twin-spool engine. Work done + work wasted in exhaust 37. Performance of the jet engine is not only concerned with the thrust produced, but also with the efficient conversion of the heat energy of the fuel into kinetic energy, as represented by the jet velocity, and the best use of this velocity to propel the aircraft forward, i.e. the efficiency of the propulsive system.
41. This latter equation can also be used for the choked nozzle condition by using vj to represent the jet velocity when fully expanded to atmospheric pressure, thereby dispensing with the nozzle pressure term (P-P9)A. Propeller efficiency at these values of V is approxi- mately 82 and 55'per cent, respectively, and from
techniques i.e., tungsten inert gas welding, resistance welding, electron beam welding and high temperature brazing in vacuum furnaces. some give resistance to corrosion whilst others can be used to release unwanted stress.
Fig. 22-3 Method of producing an engine component by sand casting.
18. Annular combustion rings can be cold forged to exacting tolerances and surfaces which alleviates the need for further machining before welding together to produce the combustion casing.
28. The most common form of tungsten inert gas welding, fig, 22-7, in use is the direct current straight polarity i.e., electrode negative pole. This is widely used and the most economical method of producing high quality welds for the range of high strength/high temperature materials used in gas turbine engines. For this class of work, high purity argon shielding gas is fed to both sides of the weld and the welding torch nozzle is fitted with a gas lens to ensure maximum efficiency for shielding gas coverage. A consumable
four per cent thoriated tungsten electrode, together with a suitable non-contact method o! arc starting is used and the weld current is reduced in a controlled manner at the end of each weld to prevent the formation of finishing cracks. All welds are visually and penetrant inspected and in addition, welds associated with rotating parts i.e., compressor and/or turbine are radiologically examined to Quality Acceptance Standards. During welding operations and to aid in the control of distortion and shrinkage the use of an expanding fixture is recommended and, whenever possible, mechanised welding employed together with the pulsed arc technique is preferred. A typical T.I.G. welding operation is illustrated in fig. 22-8.
from the actual centre of the explosion, the metal is torn into fragments which may themselves be melted by the intense heat. The dielectric medium, usually paraffin oil. pumped into the gap between the tool electrode and the workpiece, has the tendency to quench the explosion and to sweep away metallic vapour and molten particles.
ig. 22-14 Some composite material applications. 47. During the process of manufacture, component parts need to be inspected to ensure defect free engines are produced. Using automated machinery and automated inspection, dimensional accuracy is maintained by using multi-directional applied probes that record sizes and position of features. The C.N.C. inspection machine can inspect families of components at pre-determined allotted intervals 45. Composite materials have and will continue to replace casings which, in previous engines, would have been produced in steels or titanium. By-pass duct assemblies comprising of three casings are currently being produced up to 4ft-7in. in diameter and 2ft-Oin in length using pre-cured composite materials for the casing fabric. Flanges and mounting bosses are added during the manufacturing process, which are then drilled for both location and machined
without further operator intervention. In the chip machining (i.e., turning, boring, milling etc.) and metal forming processes C.N.C. machine tools enable consistency of manufacture which can be sta- tistically inspected i.e., one in ten. Component integrity is achieved by use of ultrasonic, radiologi- cal, magnetic particle and penetrant inspection techniques, as well as electrolytic and acid etching to ensure all material properties are maintained to both laboratory and quality acceptance standards.
ig. 23-2 Fuselage - mounted pod installation.
Fig. 23-1 Wing-mounted pod installation.
8. The pitot-type intake can be used for engines that are mounted in pods or in the wings, although the latter sometimes require a departure from the circular cross- section because of the wing thickness (fig. 23-5). 7. The ideal air intake for a turbo-jet engine fitted to an aircraft flying at subsonic or low supersonic speeds, is a short, pitot-type circular intake (fig. 23- 4). This type of intake makes the fullest use of the ram effect on the air due to forward speed, and suffers the minimum loss of ram pressure with changes of aircraft attitude. However, as sonic speed
9. Single engined aircraft sometimes use a pilot- type intake; however, because this generally involves the use of a long duct ahead of the compressor, a divided type of intake on each side of the fuselage is often used (fig. 23-6). 10. The disadvantage of the divided type of air intake is that when the aircraft yaws, a loss of ram pressure occurs on one side of the intake, as shown in fig. 23-7, causing an uneven distribution of airflow into the compressor. Fig. 23-7 Loss of ram pressure in divided intakes.
12. As aircraft speed increases still further, so also does the intake compression ratio and, at high Mach numbers, it is necessary to have an air intake that has a variable throat area and spill valves to accommodate
13. The engine is mounted in the aircraft in a manner that allows the thrust forces developed by the engine to be transmitted to the aircraft main structure, in addition to supporting the engine weight and carrying any flight loads. Because of the wide variations in the temperature of the engine casings, the engine is mounted so that the casings can expand freely in both a longitudinal and a radial direction. Types of engine mountings, however, vary to suit the particular installation requirement. Turbo- jet engines are usually either side mounted or underslung as illustrated in fig. 23-11. Turbo- propeller engines are mounted forward on a tubular framework as illustrated in fig. 23-13.
Fig. 23-10 Fuselage intakes. Power plant installation
14. The jet pipe is normally attached to the rear of the engine and supported by the engine mountings. In some installations, particularly where long jet pipes are employed, an additional mounting is provided, usually in the form of small rollers attached to each side of the jet pipe. The rollers locate in airframe-mounted channels and support the weight of the jet pipe, whilst still allowing it to freely expand in a longitudinal direction.
Fig. 24-2 Inspection of H.P. nozzle guide vanes.
elimination checks should normally be undertaken before more involved tasks. The manufacturers' maintenance manual contains trouble shooting information, usually in chart form and fig. 24-4 shows a typical example. 25. The reported symptom will frequently originate from flight deck instrument readings and, unless it is apparent from supporting information that the readings are genuine, instrumentation should be checked before proceeding further. Similarly, quick
Fig. 24-5 Typical friction locked adjusters.
Fig. 24-7 Ground running danger zones.
36. The starting drill varies between different aircraft types and a starting check procedure is normally used. Generally, all non-essential systems 37. At a predetermined point during the starting cycle, the high pressure fuel shut-off valve (cock) is opened to allow fuel to pass to the fuel spray nozzles, this point varying with aircraft and engine type; on some installations the shut-off valve may be opened before the starting cycle is initiated. During the engine light-up period and subsequent accelera- tion to idling speed, the engine exhaust gas temperature must be carefully monitored to ensure
38. When a turbo-propeller engine is being started, the propeller must be set to the correct starting pitch, as recommended by the engine manufacturer. To provide the minimum resistance to turning and thus prevent an excessive exhaust gas temperature occurring during the starting cycle, some propellers have a special fine pitch setting. 40. Before an engine is stopped, it should normally be allowed to run for a short period at idling speed to ensure gradual cooling of the turbine assembly. The only action required to stop the engine is the closing
Fig. 25-1 Example of growth of time between overhaul (T.B.O.).
29. Chromic acid anodizing may be used as a means of crack detection on aluminium parts, e.g. compressor blades. This process, in addition to providing an oxide film that protects against corrosion, gives a surface that reveals even the smallest flaws.
Fig. 25-4 Cracks revealed by magnetic crack detection.
Fig. 25-5 Unbalance couples due to centrifugal force.
Fig. 25-6 Dynamic balancing machine.
50. Modular assembled engines demand different balancing methods which include the use of simulated engine rotors. The dummy rotors must reproduce the bearing span, weight, centre of gravity and dynamic characteristics of the sub-assembly it
52. The engine can be built in the vertical or horizontal position, using the ram or stand illustrated in fig. 25-TO and 25-11 respectively. Assembling of the engine sub-assemblies or modules is done in separate areas, thus minimizing the build time on the build rams or stands. 51. With the introduction of the large fan blade, moment weighing of blades has assumed a greater significance, ref. fig. 25-8. This operation takes into account the mass of each blade and also the position of its centre of gravity relative to the centre line of the disc into which the blade is assembled. The mechanical system of blade moment weighing may be integrated with a computer, ref. fig, 25-9, which will automatically optimise the blade distribution. The moment weight of a blade in units i.e. g.mm. or 0z.in., is identical to the unbalance effect of the blade when installed into a disc. The recorded measurement of blade moment weights enables each blade to be
Fig. 25-9 Integrated blade moment weighing.
Fig. 21-10 International Standard Atmosphere.
_24-1 Atypical maintenance schedule (extracts).
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Mechanics of Flight Kermode
by Francisco Dávalos
The lasting popularity of this classic book is aptly demonstrated by the fact that this is the eleventh edition. This is also the third time that the current reviewers have undertaken the task of updating it, and we hope that the changes... more
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Megaproject Planning and Management: Essential Readings, vols. 1-2
by Bent Flyvbjerg

2014

Megaproject Planning and Management: Essential Readings contains the seminal articles from the growing body of research on megaproject planning and management along with an original introduction by the editor, Bent Flyvbjerg. The leading... more
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Dava J. Newman - Interactive Aerospace Engineering and Design
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Figure 5.16 | FBD with solutions to the equations of equilibrium.
Figure 5.22 | The Method of Sections cut for the truss solution.
Figure 1.1 | Leonardo da Vinci quotation and multimedia ornithopter tutorial on the accompanying CD-ROM.
Figure 1.2 | The first artificial satellite in soace, Sputnik. 27-year-old Gagarin completed one full orbit of the earth in his historic flight, which took 108 min and reached a maximum altitude of 200 km. Three weeks later, the United States countered by sending Alan Shepard into space but only on a ballistic trajectory lasting 15 min. The first U.S. orbital flight was made by John Glenn on 20 February 1962.
Figure 1.3 | The Mars rover Sojourner and the Mars Pathfinder spacecraft.
Figure 2.1 | Decavitator human-powered hydrofoil in action.
Figure 2.2 | Decavitator human-powered hydrofoil and the Boston skyline.
Figure 3.1 shows a simple wing geometry, similar to that found on a light gen- eral aviation aircraft, as viewed from above (top view). The front of the wing, at the left, is called the leading edge; the back of the wing is called the trailing edge. The distance from the leading to trailing edges is called the chord c. The ends of the wing are called the wing tips, and the distance from one wing tip to the other is called the span b. The shape of the wing, when viewed from above looking down onto the wing, is called the planform. In this figure, the planform is a rectangle, and for a rectangular wing, the chord length at every location along the span is equal. For most other planforms, the chord length varies along the span. The wing area is the projected area of the planform and is bounded by the leading and trailing edges and the wing tips. The total surface area S includes both upper and lower surfaces; the wing area is a projected area and is almost one-half of the total surface area.
Figure 3.2 | Airfoil section with lift and drag forces depicted through the quarter chord. The lift (and drag) of a given airfoil shape can be complex to calculate due to the multiple factors involved, such as airfoil geometry, flow speed, and flow type. Generally, all these parameters are lumped into an empirical nondimen- sional parameter to make the evaluation of the lift and drag more efficient. The nondimensional lift and drag terms are called the coefficient of lift and drag, respectively. [See Equation (3.10) and Equation (3.11).] The es Teh AAA As ek? 2
The velocity vector is defined to be everywhere tangent to the streamline, and the direction of ds is everywhere parallel to the local velocity, which yields where the volume of the fluid particle is an infinitesimal streamwise distance ds, multiplied by the area of the perpendicular face dA. Therefore, the mass of the fluid particle (assuming a cubic geometric shape for the fluid particle) can be written as
This can be rewritten as Bernoulli’s equation, after another 18th-century Swiss mathematician, Daniel Bernoulli:
If the shape of the blunt body is such that it allows the flow to return smoothly to a direction parallel to the incoming flow, the local pressure down- stream on the body will reestablish itself (almost) at the original upstream local static value P. In fact, this is the motivation behind the design of streamline shapes.
When a@ < a@,,,, the C,(@) (pronounced “C, of alpha’) curve depends only on the wing geometry (shape). The wing size, airspeed, and air density are irrel- evant. However, Cp is approximately the initial drag Cp 9 plus a constant k times a’. This plot depends not only on wing geometry but also on surface smoothness and the air viscosity pw. under its weight. Stall is influenced by the smoothness of the surface of the wing and the viscosity of the fluid (i.e., air viscosity for airplane flight). You can see that C, and @ have a rather linear relation up until the stall angle of attack; therefore, we can state that C, is approximately equal to ma, where m is the slope of the C,-versus-a plot.
Figure 3.7 | Surface pressures on an airfoil.
Figure 3.9 | The effect of smooth and dimpled surfaces on golf ball design.
Recall, induced drag is the form of drag on an aircraft that is related to lift. Specifically, induced drag arises from the three-dimensional effects of a wing caused by downwash velocity near the wing tip. The circulatory motion of the air at the trailing edge of the wing (called a vortex) tends to drag the surround- ing air down with it, causing a downward velocity component at the wing (see Figure 3.11). This downward component is called the downwash velocity and is often visually apparent when an airplane emerges out of a cloud bank.
Figure 3.10 | Flow separation regimes over a flat plate as determined by Reynolds number showing lamina flow transitioning over to turbulent flow.
Figure 3.12 | Lift curve for symmetrical and cambered airfoils. Figure 3.5 gave the overall shape of coefficient of lift and coefficient of drag curves as a function of airfoil angle of attack. The flow field around a cambered airfoil is that lift produced at a zero angle of attack. This is because the camber of the foil results in faster flow velocities and lower pressures above the top sur- face of the airfoil. The effect of a cambered airfoil compared to a symmetric air- foil is that the cambered airfoil’s lift curve remains about that of the symmetric foil. The cambered airfoil has the unwanted effect of flow separation occurring sooner than that in the symmetric foil; therefore, cambered foils stall at a lower angle of attack than symmetric foils. A figure that plots coefficient of drag ver- sus coefficient of lift (rather than angle of attack) is known as a drag polar. The drag polar helps show the relation of lift and drag. A cambered airfoil has mini- mum drag at a nonzero value of C, (see Figure 3.12 and Figure 3.13). The effect of Reynolds number on the lift and drag curves is shown in Figure 3.14. Higher Reynolds numbers reveal that the transition from laminar to turbulent flow occurs closer to the leading edge of the airfoil. Higher Reynolds numbers have more skin friction drag, but less pressure drag since flow separation is delayed.
The engine or engines of an aircraft produce a force called thrust that pro- pels the vehicle, as detailed in Chapter 6, “Aircraft Propulsion.” Aircraft are equipped with two stabilizers at their tail whose task is to provide stability for the aircraft and to keep it flying straight. The fixed horizontal component is
Figure 4.2 | Depiction of aircraft pitch, roll, and yaw by moving control surfaces.
Figure 4.3 | Forces acting on an aircraft in a flight inclined at a flight path angle theta to the horizon. The state variables are the velocity v, the flight path angle 0, the horizontal position x, and the altitude 4. Only one of these four requires some explanation. The flight path angle (sometimes also called the pitch angle) @ is the angle that the velocity vector v makes with the horizon (or the x axis). It is not the angle between the body axis and the horizon. The control of the aircraft consists of the magnitude of the thrust vector T and the angle of attack a. In our model we will assume that T lies in the direction of flight, which in general depends on the location of the engine(s). The angle of attack is controlled by the elevators, which is the only one of the three control surfaces coming into play in a 2-D model. The weight of the aircraft is one of the constants.
Figure 4.4 | Thrust-velocity simulation for a Boeing 727 aircraft. For a propeller-driven aircraft, we assume that the thrust available decreases approximately linearly with increasing speed. For the jet engine example above, we assumed the engines put out thrust available over the entire operating range of aircraft velocities, which is reasonable for speeds less than Mach |. However, thrust produced by propellers is significantly affected by the aircraft’s operating velocity and altitude. There is a crossover velocity where diminishing thrust available and increasing thrust required intersect, as seen in Figure 4.5.
Figure 4.5 | Thrust-velocity curve for a propeller-driven aircraft.
By examining Figure 4.6, it is easy to balance the forces. We write Let us turn to a special case, gliding or unpowered flight. In this case the thrust is equal to zero; the flight path is steady, but it is not level. Figure 4.6 shows a free-body diagram of an aircraft in an equilibrium glide.
Figure 4.7 | The diagrams show a top view and a front view of an aircraft making a level turn.
Thinking of this force as a centripetal force for a rotation, we write as for the level turn. Equating Equation (4.67) with Equation (4.68) and solving for the turning radius 7 we obtain
Figure 4.9 | The diagram shows an aircraft in a “pushdown” maneuver. Pushdown Maneuver In the third and final case, the pushdown maneuver, the aircraft is initially in a straight and level flight. Then the pilot pushes on the control column (the yoke) to deflect both ailerons to create less lift. Since now W > L, the aircraft will turn downward. The maneuver is shown schematically in Figure 4.9. We follow the same procedure to determine the turn radius and the turn rate.
Figure 4.10 | \V-—ndiagram, or flight envelope, for a Cessna Citation-like aircraft. the angle of attack is increased, the wing stalls and the load factor decreases. As the velocity increases, so does the load factor to a maximum value n,,,,, fepre- sented by point B. At point B both the lift coefficient and n are at their highest possible values that can be obtained within the allowable flight envelope. The velocity corresponding to point B is known as the corner velocity (marked by the cursor on the figure) or maneuver point denoted by v" and calculated as follows:
Figure 4.11 | Expanding the V-n diagram flight envelope of an aircraft.
The summation or addition of vector quantities such as forces is straightforward. The numerical procedure is as follows, given two vectors F and G:
If two vectors are normal to each other (i.e., perpendicular to each other), rotIr eealar nradict ic varq Ear evamnile the dat nradiicet af tuzn wit vectare ic See Figure 5.3.
The dot product (or scalar product) of two vectors returns a scalar quantity. Physically the scalar product is the length of the projection of a vector onto another vector times the length of the vector on which the projection is made. The dot product is calculated as follows:
The cross product is calculated as follows: The cross product (or vector product) of two vectors results in another vector. The resultant vector is perpendicular to a plane made up by the two vectors being multiplied. The direction of the new vector is given by the right-hand rule, and the magnitude (i.e., length) of the vector is equal to the area of the parallelogram made up by the two vectors being multiplied (see Figure 5.4).
Figure 5.5 | A schematic illustrating the concept of transmissibility of forces.
5.3.2 Supports and Reaction Forces There are four types of supports in statics, namely, rollers, pins, clamps, and fric- tion. Here you can see the reaction forces caused by each:
3.3.3 Static Equilibrium Now that we have reviewed vectors and understand the possible reaction forces, we proceed to a discussion of equilibrium and finally truss analysis. Newton’s second law can be written as
associated with the cross section of the beam. Typical values of Young’s modu- lus E are given in Table 5.1. Aircraft structures continuously undergo dynamic loading profiles, which include torsional twisting throughout the airframe. The deformation of material due to shear tends to kink the material, and the measure of deformation is often denoted by the angle @. The angle is termed the shearing strain and is analagous to the case of tension and compression where up to a cer-
Figure 5.12 | Structural application of beam theory. Figure 5.11 | Material strain-stress diagram com- paring the behavior of composites with metals [33].
Figure 5.15 | Free-body diagram for the truss structure.
Figure 6.1 | Propeller schematic showing the envelope of a moving air mass where inlet velocity is Up and exit velocity is Up.
A combustion engine’s thermal efficiency is governed by thermodynamics. We can show that the thermal efficiency is less than the Carnot (ideal) efficiency.
Figure 6.3 | Jet, or gas turbine, engine schematic with station numbers assigned to components [36].
Figure 6.4 | Typical inlet designs for subsonic and supersonic jet engines [36].
Figure 6.5 | The components of the burner, or combustor, of a gas turbine engine [36]. compressor passes through the hole. Burners must be made from materials that can withstand the high temperatures of combustion. There is usually an outer casing and an inner liner. The liner is often perforated to enhance mixing of the fuel and air. The three main types of combustors are annular, with a liner inside the outer casing (peeled open in Figure 6.5); a can, which is a tubular design where each can has a liner and a casing; and a hybrid can-annular, where the cas- ing is annular and the liner is can-shaped. Many modern burners incorporate
Figure 6.7 | Gas turbine nozzles [6].
Figure 6.8 | High bypass turbofan engine schematic. Figure 6.9 | Low bypass turbofan engine schematic.
front of the engine and are driven by a fan turbine located behind the primary turbine that drives the main compressor. For supersonic flight, a low bypass fan is used that has a much smaller front fan, and often afterburners (discussed in Section 6.4.4) are added for additional thrust. Even low bypass ratio turbofans are more fuel-efficient than basic turbojets. They can then cruise efficiently but have sufficient thrust for dog-fighting and military maneuvers. Even though the fighter plane can fly much faster than the speed of sound, the air going into the engine must travel at less than the speed of sound for high efficiency. The air- plane inlet slows the air down from supersonic speeds.
Figure 6.10 | Turboprop engine schematic. The thrust of a turboprop is the sum of the thrust of the propeller and the thrust of the core. We can use our basic thrust equation on the propeller and core to obtain the thrust equation for the turboprop. As we have noted above, the mass flow through the propeller is much greater than the mass flow through the core engine. And we have also noted that the exhaust velocity of the core is low and almost equal to the velocity into the core. The mass flow rate exiting the core is almost equal to the mass into the core. Comparing with the pure propeller the- ory, the thrust is equal to the mass flow through the propeller times the velocity change across the propeller plus a smaller amount of thrust from the core engine. Because propellers become less efficient as the speed of the aircraft increases, moving a large mass of air through a small change in velocity. Propellers are very efficient and can use nearly any kind of engine (including humans!) to turn the propeller. There are two main parts to a turboprop propulsion system: the core engine and the propeller. The core is very similar to a basic turbojet except that instead of expanding all the hot exhaust gases through the nozzle to produce thrust, most of the exhaust energy is used to turn the turbine. There may be an additional turbine stage present, which is connected to a driveshaft. The shaft drives the propeller through gear connections and produces most of the thrust. The exhaust velocity of a turboprop is low and contributes a small amount of thrust since most of the energy of the core exhaust has gone into turning the drive- shaft (see Figure 6.10).
Afterburners are only used on fighter planes and the supersonic Concorde airliner. (The Concorde turns the afterburners off once it gets into cruise. Other- wise, it would run out of fuel somewhere over the Atlantic.) Afterburners offer a mechanically simple way to augment thrust and are used on both turbojets and turbofans. overcome a sharp rise in drag near the speed of sound. A simple way to get the necessary thrust is to add an afterburner to a core turbojet, often called an aug- mentor. In a basic turbojet, some of the energy of the exhaust from the burner is used to turn the turbine. The afterburner is essentially a long tailpipe into which additional fuel is sprayed directly into the hot exhaust and burned to provide extra thrust. Figure 6.11 shows that the nozzle of the basic turbojet has been extended, and there are now a couple of flame holders in the nozzle. In three dimensions, the afterburner is composed of a series of hoops placed in the exhaust stream. When the afterburner is turned on, additional fuel is injected through the hoops and into the hot exhaust stream of the turbojet. The fuel burns and produces additional thrust, but is less efficient than in the combustion sec- tion of the turbojet. The additional thrust comes at the cost of much more fuel. When the afterburner is turned off, the engine performs as a basic turbojet. The mathematics of an afterburning jet is the same as that for a nonafterburning jet as far as thrust is concerned. The exhaust velocity is increased compared to that of the nonafterburning jet because of the higher temperatures involved.
A rocket in its simplest form is a chamber enclosing a gas under pressure. A small opening at one end of the chamber allows the gas to escape, and in doing so it provides a thrust that propels the rocket in the opposite direction. Think of the example of a balloon. Air inside a balloon is compressed by the balloon’s rubber walls. The air pushes back so that the forces on each side are balanced. When the nozzle is released, air escapes through it and the balloon is propelled in the opposite direction. For spacecraft, the gas is produced by burning propel- lants that can be solid or liquid in form or a combination of the two.
Figure 6.15 | Screen of the turbofan engine performance simulator.
Figure 7.1 | Definition of the airplane’s axis system. A classic airplane has three basic controls: ailerons, elevator, and rudder. They are designed to change and control the moments about the roll, pitch, and yaw axes. These control surfaces are flaplike surfaces that can be deflected back and forth at the command of the pilot.
Figure 7.2 | Dynamically stable behavior. In both situations the airplane eventually returns to its equilibrium position after some time interval. In the first case the vehicle approaches monotonically the equilibrium position (aperiodic behavior), while in the second case, it over- shoots the equilibrium position (damped oscillation). A body is dynamically sta- ble if, out of its own accord, it eventually returns to and remains at its equilib- rium position over a period of time. It is important to note that static stability does not imply dynamic stability, as Figure 7.3 shows. The plane is dynamically unstable but still statically stable.
1. Pressure is the force perpendicular to the surface per unit area, while shear stress is the force along the surface per unit area. Both have units of pascals (or newtons per meter squared).
Figure 7.3 | Dynamically unstable behavior. 7.3 | STATIC FORCES AND MOMENTS ON AN AIRCRAFT 7.3.1 Resulting Force on a Wing
Figure 7.5 | Contributions to the moment acting about the center of gravity. Having looked at a wing only, we can now consider a complete airplane, as shown in Figure 7.5. In examining a whole aircraft, the pitching moment about the center of gravity (center of mass) is of interest. The moment coefficient about cg is defined analogous to the moment coefficient about the ac:
Figure 7.6 | The moment coefficient about the center of gravity as a function of angle of attack for a longitudinally stable aircraft. If the airplane is flying at its trim angle of attack a, and suddenly encoun- ters a disturbance that causes it to pitch up or down (e.g., due to a wind gust), the moment will be such that the plane will return to its equilibrium position. To see that, imagine a wind gust pitching the plane up from a, to some larger a. By looking at the plot in Fig. 7.6, you can see that the moment coefficient (and hence the moment) will be negative, which makes the plane pitch down and return to equilibrium.
Figure 7.7 | The moment coefficient about the center of gravity as a function of angle of attack for a longitudinally unstable aircraft. That is, the slope of the moment coefficient curve versus angle of attack has to be negative, and the moment coefficient at zero angle of attack has to be positive.
Figure 7.8 | Forces acting on the tail and relative wind seen by the tail [33].
Figure 7.9 | Geometry of a wing-tail combination [39].
Figure 8.1 | The first impression one has looking back at Earth from Low Earth Orbi (LEO) is the realization that our atmosphere is a delicate, almost invisible, shield. Breathable air, about three human hairs thick on our basketball model, is a powerful shield against more than meteors. It has a complex chemical structure. It is mostly oxygen and nitrogen at sea level, but it becomes richer in carbon dioxide (CO,) and ozone, at altitude. Standard sea-level pressure and tempera- ture are 101.3 kPa and 290 K (17°C), respectively. The density of air on Earth is equal to 107° atoms per cubic meter (1.22 kg/m), which is considered to be a standard 1 atmosphere of pressure (atm, or 101.3 kPa).
Figure 8.2 | Variations in pressure and temperature as altitude increases. (from [50], courtesy of NASA.) 1. The troposphere begins at sea level and reaches a height of about 18 km. The temperature in this region varies with altitude from about 290 K (17°C) at sea level to 220 K (—53°C) at 11 km. Nearly all the weather effects we experi- ence on the surface of Earth, such as rain and snow, occur within the tropo- sphere. As is shown in the temperature gradient, the surface of Earth is the warmest part of the troposphere. Heat is transferred upward by means of infrared
3.2.2 The Temperature Extremes of Space Figure 8.3 | Atmospheric atomic oxygen density in low Earth orbit.
Figure 8.4 | Mir21/NASA2—Mission photo taken of the Russian Mir space station. The first impression one has looking back at Earth from Low Earth Orbit (LEO) is the realization that our atmosphere is a delicate, almost invisible, shield.
Figure 8.7 | Van Allen inner and outer belts. [41] Two concentric regions of trapped electrons have been identified: The first extends to about 2.4 Earth radii, and its radiation particle population is rather sta- ble; the second is an outer belt from 4 to 6 Earth radii whose electron density is highly variable and very sensitive to solar activity. The doughnut-shaped Van Allen belts are distributed nonuniformly within the magnetosphere, and extended stays in either could be fatal to humans [55]-[57].
Figure 8.8 | Earth’s magnetosphere illustrating key energetic particles that respond to variations in solar wind parameters. [41]
Figure 8.9 | Orbital debris population distribution. The aesthetic feature of this space debris picture raises our concern and con- sciousness of the problem, and the dangers posed by these human-made debris to operational spacecraft (pilotless or piloted) are a growing concern. A dramatic illustration of this fact is the recent impact of the mini satellite Cerise with an Ariane 4 third stage. Ground controllers lost contact with Cerise, a small British- built satellite, only to discover weeks later that the satellite was struck by a frag- ment from an Ariane rocket which had exploded several years previous. The col- lision broke off part of the satellite that helped maintain its stability. During a recent Hubble Space Telescope repair mission, astronauts found a small hole in
Figure 9.1 | Geocentric universe is replaced by the Copernican theory of < heliocentric universe. Quantitatively, at least, the architecture for the universe devised by Coper- nicus accounted for several of the mechanisms observed in the sky: It explains the rising and setting of the Sun, Moon, and stars; and the aggregate motion of
Figure 9.2 | Conic sections: circle, ellipse, parabola, and hyperbola.
Figure 9.3 | Focus-Directrix geometric property of an ellipse. en ee OEE IE EEE EID EI EI Ellipses demonstrate three basic and well-known geometric properties: the focus-directrix property, the focal radii property, and the orbital tangents prop- erty (see Figure 9.3). The focus-directrix property states that the orbit is the locus of points whose distances from a fixed point and a fixed straight line are in constant ratio. The focal radii property declares that an ellipse is the locus of points for which the sum of the focal radii is constant. The focal radii are the dis- tances to a point P on the ellipse from the two foci of the ellipse. The property of orbital tangents states that the focal radii to a point on an ellipse make equal angles with the orbital tangent to that point of the ellipse (see Figure 9.4). In the usual systematic development of analytic geometry, one of these properties is generally chosen as the definition of this special class of plane curves.
A direct consequence of the law of areas is Kepler’s equation of orbit, which relates the orbital position with time. But, first, some nomenclature must be expounded upon. An elliptical orbit of a body can be characterized by its semi- major axis and its eccentricity. The semimajor axis a of an ellipse can be described as one-half the length of the longitudinal axis of the ellipse. The eccen- tricity e defines the shape of the orbit; it is quantified as the apogee radius (max- imum) minus the perigee radius (minimum), divided by their sum, as defined in Equation (9.11). Figure 9.5 illustrates these parameters of an elliptical orbit.
Figure 9.5 | Illustration of Kepler’s third law and geometry of an The eccentric anomaly can be constructed by circumscribing a circle around the elliptical orbit of the body; this circle is tangent to the two ends of the ellipse (at perigee and apogee for an earth orbit). A perpendicular from the semimajor axis is erected from point R, passing through the orbital position, P, of the body on the ellipse and intersecting the great circle at Q. A line is then drawn joining the perpendicular intersection point on the circle to the center of the ellipse, C, (and thus the center of the circle) (see Figure 9.6). The eccentric anomaly, E, is measured from the perigee point in the direction of travel of the satellite to the line joining the center of the ellipse to the intersection point. The mean anomaly, M, of a body for a given position and time is the angular position subtended by a fictitious body traveling for the same amount of time on a circular orbit which possesses the same period as the original body’s elliptical orbit constant rate. The mean anomaly and eccentric anomaly are related through Kepler’s equation:
iterative methods. Hence, the corresponding radial distance, 7, and true anomaly, f, can be calculated from Equation (9.14) and Equation (9.15).
Hence, for any given transfer orbit, the velocities can be calculated graphi- illy if the initial radius vector, final radius vector, and transfer angle are known. result occurs if a change of coordinate system is employed. When the terminal velocity vectors v, and v, are resolved into components parallel to the chord connecting the endpoints and parallel to their respective radius vectors (that is, v, is resolved parallel to c and to r,), the magnitudes of components of the veloc- ity vectors along these skewed axes at the two terminal points are, respectively, equal. This means that the chordal and radial components of the terminal veloc- ities of any given position vectors are always equal (see Figure 9.7).
Figure 10.1 | Satellite mass as represented by percentage of the overall design [41].
The antenna requires a signal that is powerful enough to be recognized by the electronics that decode the signal. Thus orbital altitude can be used to deter- mine a preliminary power requirement for sizing the communications system (i.e., set the altitude, determine the minimum signal strength that can be received, and calculate the required power for the satellite transmitter). Figure 10.2 | Schematic of communication and telemetry data flow paths showing uplink and downlink.
form (ignoring transmission losses and signal attenuation through the atmo- sphere), the communications link power budget is given by
Figure 10.4 | The four primary heat sources for satellites: direct solar heating, reflected solar energy from the Earth’s albedo, the Earth’s infrared energy, and on- board sources. The only way to dissipate heat from a satellite is by radiating heat away from the satellite, as there is no atmosphere in space for convective cooling. On board the satellite, heat is transported via conduction between elements, heat pipes that use a working fluid, and internal radiation. The thermal designer must utilize arrangements of materials, coatings, and heat conduits on board satellites to main- tain the components within their operating limits. In some cases heaters must also be added to maintain temperatures within acceptable limits. An example of tem- perature ranges to be maintained for satellite components is shown in Table 10.2.
Figure 10.5 | Representative material selection for satellite components.
Figure 10.6 | Attitude determination and control subsystem block diagram. Five major references are typically used in satellite design, including the Sun, the Earth, Earth’s magnetic field, stars and planets, and inertial sensors (gyro- scopes). The Sun is undoubtedly a bright and unambiguous light source. Optical Sun sensors require relatively little power and are lightweight. However, satellites in LEO can lose the Sun reference if eclipsed. Additionally, the Sun only provides measurement about two independent axes. Earth is an excellent reference for Earth-orbiting satellites as it is close and bright. Earth sensors scan to detect the horizon, the boundary between the blackness of space and the brightness of Earth. Earth sensors must be protected from sunlight as they are more sensitive.
Figure 10.7 | Bi-propellant rocket nozzle schematic. For ideal situations, the thrust can be maximized by expanding the exiting gas(es) fully to the ambient pressure of the surrounding environment. For first- order estimates, this reduces Equation (10.4) to
Figure 11.1 | Earth as seen from the Moon by the crew of the Apollo 17 mission.
Figure 11.2 | Skylab demonstra- tion of the microgravity environment, or weightlessness.
Figure 11.3 | Launch of Shuttle Mission STS-96, also called ISS mission 2A.1, to resupply the ISS on 27 May 1999. Between March 1995 and May 1998, NASA astronauts flew onboard the Russian space station Mir in a collaborative effort. The NASA program sup- porting this endeavor is commonly known as International Space Station Phase I (or Shuttle-Mir), encompassing 11 Space Shuttle flights and joint Soyuz flights. The international program resulted in joint space experience for the crew and the start of joint scientific research. Shuttle-Mir participants (crew members, princi- pal investigators, mission control staff) investigated vital questions about the future of human life in space. Mir is a test site for three main areas of experience and investigation:
Figure 11.4 | The ISS fully assembled after mission 16A (“A” denotes an American launch of the final component, the U.S. habitation module, shaded in drawing). Figure 11.5 | The “Wall of EVA” illustrating the history of EVA and the three-fold increase in EVAs required for ISS assembly.
Figure 11.7 | Cylindrical fabric soacesuit component upright (left) and deformed joint (right). where 6 is the deformation joint angle. Substituting into Eqn. (11.2) yields
Figure 11.8 | Spacesuit elbow joint with constant centerline length and flexed spacesuit joint. The force trying to elongate the joint due to spacesuit internal pressure F= PA.
Figure 11.9 | Possible EVA requirements for a planetary mission where HST is Hubble Space Telescope anc DTO stands for a detailed test objective, or additional EVA opportunities. Advanced spacesuits should provide the following: working pressures that allow shirtsleeve mobility, dexterous gloves or actuators, longevity, ease of maintenance, and adequate environmental protection. Mobile joint systems must allow for minimum energy expenditures during EVA tasks. Gloves should be certified for long-duration use at high pressures. Improved technology and mate- rials should allow for durable spacesuit design. Advanced reliable primary life support systems should be regenerable, low-mass, and modular. A broad meta- bolic loading range between 63 and 625 kcal/h (250 and 2500 Btu/h) should be realized with automatic thermal control systems [124]. A modular, evolvable design is advantageous. Technological advances should lead to real-time envi- ronmental monitoring systems and innovative display and vision systems. A Martian EVA will most likely entail rappelling down shear cliffs and traversing monumental canyons. A radical new approach to spacesuit design could best realize these challenges.
Heel-off interactions with knee Figure 11.11 | Top. Heel strike. The foot plantar flexes which lowers the ankle as the foot contacts the ground. Bottom. Heel-off interactions with the knee. Heel-off keeps the excursion of the center of gravity to a minimum.
characteristic of gait. At heel strike the foot plantar flexes (rotates downward about an axis formed at heel contact), thus lowering the ankle as the foot makes full contact with the ground (see Figure 11.11). A fused ankle joint (immobile) without plantar flexion would cause the cg to rise as if the leg were a stilt. Ankle plantar flexion affects gait similarly to ankle flexion in that the trajectory of the cg is reduced and shock absorption is noted at heel strike. The heel-off phase pro- vides a horizontal cg trajectory as the ankle rotates upward about an axis formed at the ball of the foot. The trunk flexes both laterally and anteroposteriorly dur- ing walking to make up the final characteristics of walking. The ipsilateral flex- ion of the vertebral column toward the stance-phase side causes a 1 to 2 cm dis- placement. The anteroposterior flexion of the trunk reveals maximum backward flexion at the beginning of the support phase and maximum forward flexion toward the end of the support phase, resulting in small 1 to 2 cm deflections. Figure 11.10 | Pelvic rotation during walking. The pelvis is rotated from side-to-side about the longitudinal axis of the body.
*Stone, R. W. (1971) Man in Space. Figure 11.13 | Stepping frequency for Apollo 11 and simulated lunar gravity. Stepping frequency for terrestrial locomotion is also plotted. The Apollo data and simulated lunar data show a reduction in stepping frequency as compared to 1 G, especially for locomotion at velocities of 1.5 m/s and 2.3 m/s [126].
Figure 11.12 | Locomotion force traces at 1 g, and simulated Martian gravity and lunar gravity showing a significant reduction in peak force (fina,) for each stride during Martian and lunar locomotion as compared to G locomotion. The time of foot ground contact is denoted by t,, time of flight (or aerial time, t,), and the time for the entire stride, toride-
Figure 11.14 | Oxygen requirements for space flight. On all U.S. missions to date, the O, has been supplied from tanks that car- ried sufficient O, for the duration of the mission. For longer-duration missions, the storage or resupply penalty is excessive, and some method of recovering O, from waste mass is required. For the International Space Station, the O, will ini- tially be resupplied, but will later be generated by electrolysis of recovered wastewater. Other methods include electrolysis of water vapor and electrolysis of CO,. There are also biological methods for generating O,, including algae or other microorganisms and higher plants such as salad vegetables [132].
Figure 12.1 | Hands of Creation.
Figure 12.2 | A parallel bar, triangle, scale, template for drawing circles, and electric eraser are shown on a drafting table used for drawing. 3. Pencils. Once a method of moving around the drawing plane is estab- lished, the next concern is with marking the drawing plane. Many tools can be used to do this, but in designing, the ability to erase is very important. A well- sharpened, clean pencil is best to make sharp lines of varying weights (discussed later) while allowing the designer to erase at any time.
Figure 12.3 | An example of the proper technique for drawing straigh lines. Make sure the point of the pencil is always sharp and clean. After sharpen- ing the lead, wipe the tip with a tissue to remove excess graphite. Then, with a scrap of paper slightly dull the tip of the lead so that it does not break when the tip is placed to the paper. Breaking the pencil tip will not only cause frustration, but also make an uneven line and cause graphite shavings to smear the drawing. To keep the point sharp while drawing, the pencil can be rotated in the hand. This technique takes some getting used to, but it ensures that the lead is worn down evenly, preserving the point.
Figure 12.4 | An artist's conception of a space station.
Figure 12.7 | Illustration of one-point perspective. Two-Point Perspective: A two-point perspective drawing contains two vanishing points. It is used when a principal plane of the object is not parallel to the picture plane; however, a set of parallel lines on the object may remain parallel to the picture plane. For example, imagine looking at the corner of a building or at the intersection of two streets (the lines of each street converge to their own vanishing point) (see Figure 12.8).
Figure 12.8 | Illustration of two-point perspective.
Figure 12.9 | Illustration of three-point perspective.
Figure 12.10 | Lighter-than-air (LTA) vehicle race course around a gymnasium.
Figure 12.11 | Dimensioned LTA vehicle drawing from the CDR.
Figure 12.12 | Examples of prototype pitch control and rendered LTA vehicle presented during a CDR.
Figure 12.13 | Personal design portfolio examples.
Figure 12.15 | Using vectored thrust to pitch an airship (see animation in LTA Interactive on the accompanying CD-ROM). Stability An unstable LTA vehicle will require excessive control inputs to make it fly a desired course; therefore, a stable aircraft will be faster and easier to fly. A good test for stability is whether the vehicle will fly straight and level with no control inputs and will stop turning when a control input is discontinued. In order for a craft to be stable, the LTA vehicle must be well balanced, and the motors must return to their neutral position when the control input is stopped. One method for increasing stability is to use stabilizers. There are two main configurations to consider.
Figure 12.16 | Vertical and horizontal stabilizers on an LTA vehicle. Stabilizers will add to the stability about all three axes (roll, pitch, and yaw). See Figure 12.16. These stabilizers use air resistance to steady the craft. The air resistance is an effective way to counterbalance the vectored thrust forces pro- duced by the motor and propeller. When an input force is applied to turn the craft, air creates an equal opposing force against the stabilizer. When the turning force is removed, air resistance against the stabilizer slows the turning motion, returning it to straight and level flight. There are many different ways stabilizers can be incorporated into the LTA vehicle design.
Figure 12.17 | LTA vehicles take flight during the race competition. final design around the marked course. This is a time of glorious victory even for the teams whose LTA vehicles do not place in the competition; the ultimate vic- tory is the realization of the design process and for the first time knowing what it is to be a designer. This initial excitement of design from the LTA experience is carried over to a fulfilling career in engineering.
12.4 Follow the instructions given in Problem 12.2.
12.3. Follow the instructions given in Problem 12.2.
Figure 12.19 | Delta Design two-dimensional game board.
Figure A.1 | Definition of the plane angle SI unit, radian, and the solid angle SI unit, steradian. Table A.2 | Derived units with special names
Table A.5 provides conversion factors for common units. To convert temperatures between the Celsius, Fahrenheit, Kelvin, and Rankine scales, use the following formulas: A.3 | UNIT CONVERSION FACTORS Table A.5 | Unit conversion factors
Table 4.3 | Designing for a large mass ratio Let us investigate the first two design parameters further, starting with the mass ratio. A large m,,,)/9 can be obtained through proper selection of aircraft structures, materials, and overall design, as shown in Table 4.3.
Table 4.4 | Designing for a large L/D ratio
Table 4.2 | Aircraft range comparison To further understand aircraft range and the guiding performance parame- ters, let us compare the Boeing 747 four-engine jet aircraft and the Voyager around-the-world propeller aircraft. To attain the maximum range, four tradeoffs in Table 4.2 are highlighted.
Table 5.1 | Typical material properties (adapted from [34]) Figure 5.13 | Structural application of torsion.
Table 8.1 | Calculated mean temperatures of the outer planets Aluminum Satellite Considerations where T, is the average temperature of the Sun and R, is the radius of the Sun. Analyzing the equations for absorbed and emitted power, we find that the equi- librium temperature not only depends upon the distance of the body from the Sun, but also varies directly with the value of a/e. Estimating the mean temperature of Earth, using a = 0.6 and ¢ = 0.9, yields a value of 257 K. The mean tempera- tures of the outer planets, calculated in a similar manner, are found in Table 8.1.
Table 8.2 | Solar Absorptivity and Infrared Emissivity of Selected Materials
Table 10.1 | Satellite communication frequency bands (comsat=communications satellite) [41] The communications frequency bands used on satellites have been devel- oped by the International Telecommunications Union (ITU) and the World Admin- istrative Radio Conference (WARC). In the United States, the Federal Commu- nications Commission (FCC) administers the commercial frequencies, while the Interdepartmental Radio Advisory Committee administers frequencies for mili- tary uses. Table 10.1 provides a list of the allocated communication frequency
Table 10.2 | Design temperature ranges for satellite components [41] As discussed in Chapter 8, “The Space Environment: An Engineering Per- spective,” all materials have the ability to absorb, reflect, and emit certain amounts of thermal energy. The absorptivity a and emisivity ¢ of a surface are the fractions of energy emitted or absorbed compared to a perfect black body. Materials that absorb a large amount of heat are said to have high absorptivity. Likewise, materials that have the ability to radiate a large amount of energy are said to have high emissivity. Examples of material properties used in satellite applications are given in Table 10.3.
Table 10.3 | Selected satellite material emissivities and absorptivities
Table 10.4 | Various rocket types, specific impulse, and thrust range To design the satellite’s propulsion system, you need to decide on the num- ber, type, and size of thrusters as well as the amount of fuel needed for the mis- sion. The orbit chosen for the mission helps determine the types of perturba- tions for which the thrusters need to compensate. Orbital maneuvers are typically quoted in terms of velocity changes, or Av. To calculate Av, start with Newton’s law: Table 10.4 provides an overview of some rocket types and their associated specific impulse and thrust range. For comparison purposes, the first two rocket types are for large-thrust rockets, whereas the others are for satellite thrusters. The point of interest is that a high-/sp rocket, although highly desirable from a propulsive efficiency standpoint, may only have a maximum thrust output that limits it to certain applications. As an example, ion engines offer an incredibly high /sp compared to the liquid oxygen/hydrogen main engines of the Space Shuttle. However, the ion engines can only produce thrusts of under 1 N—not very effective for putting satellites into orbit, but effective for very small satel- lite maneuvers. For satellite control and maneuvers in orbit, small attitude con- trol thrusters range from 0.1 to 50 N, while large orbit control thrusters can range from 100 to 2,500 N.
Table 10.5 | Topex/Poseidon spacecraft description The satellite was launched on August 10, 1992 on an Ariane 42P from the European Space Agency’s Guiana Space Center located in Kourou, French Guiana. It orbits Earth at an altitude of 1,336 km (830 mi) with an inclination angle of 66° and a period of 112 min, carrying two altimeters onboard built by NASA and CNES. The satellite measures the height of the ocean at the same location every 10 days using two laser altimeters. Topex/Poseidon distributes the data to 38 science investigators from nine countries. Topex/Poseidon measures the satellite altitude above the sea surface while three independent satellite track- ing systems measure the satellite’s position, that is, the distance between the satellite and the center of Earth (geocenter). Then the altimetric measurements are subtracted from the satellite position, resulting in a precise height measure- ment of the ocean above the geocenter, which is the sea level. The satellite meas- urement system is shown on the CD-ROM.
Table 10.6 | Landsat 7 mission specifications
Table 10.7 | The Landsat 7 remote-sensing data applications
Table 10.10 | Satellite design trade analysis The initial parameters of your spacecraft are as shown in Table 10.10. The increase in cost of your system is a function of two factors: (1) the increased cost of the solar array, due to using more expensive solar cells, and/or a larger array area, and (2) the cost of reducing the mass of other parts of the system if you choose to increase the solar array area. The second factor models the case where the launch mass is constrained, and so if you increase the solar array area, you must decrease the mass of other parts of the vehicle to compensate.
Table 11.1 | Characteristic astronaut motions
Table 11.2 | Average metabolic rates for past space missions [117, 118, 119]
Table 11.3 | Calculated forces for spacesuit joints There are several ways to accomplish the design of a constant-volume fab- ric component. The EMU is constructed so that as the volume of the fabric com- ponent decreases inside, it increases by the same amount outside the joint. This is accomplished by holding the centerline of the joint at constant length instead of the outside of the joint. An elbow joint designed on this principle looks like the diagram in Figure 11.8. The axial restraint lines, located across the diagonal of the fabric cylinder, take the pressure load that tries to elongate the joint. This prevents the fabric cylinder part of the joint from carrying that load and allows for excess fabric to be placed on the outside of the joint. Without the axial restraint line, the pressure would cause the joint to elongate until all the excess fabric were placed in tension. As the joint is bent with use, the inside of the joint folds up just as it did in the fabric cylinder. As that happens, the outside of the joint where the excess fabric has been placed expands to compensate for the lost volume. The flexed joint is depicted in Figure 11.8. If this figure is considered as a free-body diagram of a spacesuit joint, careful observation shows that the forces are not balanced. In practice, a spacesuit joint does not bend so that the centerline is in the shape of a circle, as shown in the figure; but rather, the cen- terline shifts slightly to the outside and compensates for the otherwise unbal- anced forces. In addition, the axial restraint lines are not placed exactly at the joints’ centerline, but they are slightly offset so that the joint will be balanced and stable. If the placement of the axial restraint line is not done carefully, either the joint is very difficult to bend, or it might actually bend itself over when pressur- ized. Therefore, well-designed joints are stable and remain where they are placed with little restraining or springback force.
Table 11.4 | Comparison between the U.S. and Russian spacesuit [123], p.43
Table 12.2 | Large-motor (210 g) performance
Table 12.3 | Small-motor (89.7 g) performance Smaller motors are also available. Data from the smaller Edmund Scientific drive motors with the propeller attached are provided in Table 12.3. The motor works approximately two-thirds as well in reverse due to propeller inefficiencies.
Table A.1 | SI base units The International System of units, referred to as SJ, is the most comprehensive and advanced system of units. SI units are divided into three categories: (1) base units, (2) supplementary units, and (3) derived units. The base units are seven well- defined units which are dimensionally independent. They are listed in Table A.1.
'Use of the prefixes in the shaded table cells is not recommended. Table A.3 | SI prefixes!
Table B.1 | Physical constants
Table B.4 | Properties of selected engineering materials
Table B.2 | Astronomical values
Table B.3 | Properties of air' 'Dry air at a temperature of 20°C and a pressure of 1 atm.
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National Academy of Engineering: Introduction to Open Hardware in Science and Engineering
by Joshua Pearce and 
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Lucas Osborn
The articles in this issue look at how the development and use of free and open source hardware (FOSH or simply “open hardware”) are changing the face of science, engineering, business, and law. Free and open source software (FOSS) has... more
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by International Journals for Researchers [ER Publication, WOAR Journals, IJEAS and IJEART]
— This investigation present a new method for reducing electrical field in power electrical substation and also it has been paid attention to health personnel who work in this place therefore the health of staff has been a priority. In... more
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Avionics Human-Machine Interfaces and Interactions for Manned and Unmanned Aircraft
by Roberto Sabatini and 
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Alessandro Gardi

2018, Progress in Aerospace Sciences

Technological advances in avionics systems and components have facilitated the introduction of progressively more integrated and automated Human-Machine Interfaces and Interactions (HMI2) on-board civil and military aircraft. A detailed... more
y. 2. Boeing 727 classic flight deck.
Historical evolution of civil flight decks. This section presents an overview of the HMI’ for manned and unmanned/remotely piloted aircraft. The evolution of civil flight decks is first presented, followed by a brief overview of military cockpits, and finally some emerging concepts for Single-
Attitude directional indicator. The Attitude Directional Indicator (A D1) as illustrated in Fig. 4 is an evolution of the attitude indicator. Besides providing basic attitude information, the A DI incorporates flight director overlays to provide instruction on intercepting and maintaining the desired flight path. To support Instrument Landing System (ILS) approaches the ADI also displays glideslope, localizer, speed deviation and decision height information. The flight director display is connected to a flight director computer, which computes the necessary guidance instructions from altitude, airspeed, attitude, heading, navigation, navigation aid (NAVAID) data (e.g., from VOR/DME) as well as autopilot mode inputs. Inclinometers, used in the Turn and Slip Indicator, are an additional component of ADI to assist the pilot in coordinating turns.
y. 5. Horizontal Situation Indicator, [10]. The Horizontal Situation Indicator (HSI), depicted in Fig. 5, is an evolution of the heading indicator. In addition to providing basic heading information, the HSI also includes a Course Deviation Indicator (CDI) overlay to support radio-based navigation. When the aircraft’s Very High Frequency (VHF) Omnidirectional Range (VOR) receiver is tuned to the frequency of a selected VOR station, the course select knob on the HSI can be tuned to intercept a chosen radial from the VOR station. The selected course is indicated on the course select pointer. The TO/FROM indicator is used to show if the aircraft is flying towards (if pointing in the same direction as the course select pointer) or away from (if pointing in the opposite direction as the course select pointer) the VOR station. The course deviation bar indicates left/right deviations from the selected course; each interval on the course deviation scale corresponds to a deviation of 2 degrees. The HSI is also used in ILS approaches, but instead of indicating VOR information, the CDI overlay now shows the course deviation on the localizer. Glideslope pointers are also displayed on the HSI during ILS approaches.
y. 7. Boeing 757, first generation flight deck, [12]. Avionics developments in the mid-1970s allowed significant digitalisation of flight deck systems, leading to the consolidation of a number of legacy interfaces into combined electronic displays. Figure 6 illustrates the historical evolution of first, second and third generation electro-optical deck displays, which included the Electronic Attitude Director Indicator (EADI) and Electronic Horizontal Situation Indicator (EHSI); Electronic Centralised Aircraft Monitor (ECAM) or Engine-Indicating and Crew-Alerting System (EICAS); as well as Multipurpose Control and Display Unit (MCDU) of the Flight Management System (FMS). The Boeing 757 (shown in Fig. 7) and Airbus A310 were among the first aircraft to feature these electronic displays.
ig. 10. | ECAM status pages for different systems, image courtesy of FAA [21].
fig. 11. | Boeing 747-400, second generation flight deck, [24].
Fig. 13. Evolution of current second generation flight decks: conceptual image from ALICIA (All Condition Operations and Innovative Cockpit Infrastructure), research and development project co- funded by European Commission under the Seventh Framework Programme [34], courtesy of Airbus and Deep Blue. Third generation flight decks (Fig. 13) are likely to see a consolidation of second generation displays and interfaces. Advances in display technology will pave the way for larger MED, providing fully integrated and interchangeable displays. Synthetic Vision Systems (SVS) are already in use in some business jets such as the Bombardier Challenger 650 and are likely to be integrated into the displays of third generation civil flight decks. EVS such as Heads-Up Displays (HUD) are being offered on some civil aircraft such as the Boeing 787 and the Airbus A320. Both SVS and EVS provide increased safety for flight in low visibility conditions such as darkness, smoke, fog or rain. SVS and EVS concepts are currently being evaluated by NASA [32, 33] for future operations.
‘ig. 12. | A CDTI display containing both plan and vertical traffic displays.
Fig. 14. | AnEVS concept using the pathway-in-the-sky representation to provide approach guidance, image used with permission of the author [37].
ig.17. _ F-15A Eagle, a fourth generation fighter, image courtesy of the National Museum of the U.S. Air Force. allows for other for other uses in the area (such as incorporating alternative control devices like keyboards, trackballs, etc.). The Gulfstream G500 is the first civil aircraft to use an active inceptor system to provide pilots with tactile feedback.
jy.16. | F-4 Phantom IJ, a third generation fighter.
j.18. | Eurofighter Typhoon, a four-and-a-half generation fighter. Reduced aircraft development due to forced reductions in military spending led to a half generation increment from fourth generation fighter jets. Four-and-a-half generation fighters such as the Eurofighter Typhoon (Fig. 18) included avionic improvements such as the Active Electronic Scanned Array (AESA) radar, integrated Infrared Search and Track (IRST) systems and high capacity data-link, allowing for network centric warfare. Four-and-a-half generation cockpits also feature altemative input technologies such as Direct Voice Input (DVI) and touchscreen displays.
Fig. 20. Embraer EMB-505 Phenom 300 flight deck, [65]. SPO flight decks are currently used in the military, General Aviation (GA) and business jet domain, but given developments in flight deck autonomy and HMI’ evolutions, SPO is expected to become a viable concept of operations for future commercial airliners. SPO flight decks contain HMI’ elements of second and third generation two-pilot civil flight decks - business jets such as the Embraer Phenom 300 (Fig. 20) are certified for SPO and feature streamlined interfaces containing wide-screen and high resolution MFD. The MFD functionalities can contain SVS, touchscreen and satellite weather capabilities that are typical of third generation civil flight decks.
‘ig. 19. F-35 Lightning II, a fifth generation fighter. Fifth generation fighters such as the F-35 (shown in Fig. 19) have completely replaced the HUD with HMD as the primary flight display. The F-35’s HMD uses a distributed aperture system to provide a 360-degree view of the aircraft's surroundings, supporting even greater situational awareness. The F-35 also features a panoramic touchscreen MFD, similar to the type of displays that will be featured in next generation of civil flight decks. Fifth generation fighters are also typically single- seated, with the role of the Weapon Systems Officer (WSO) in the back seat being replaced by other networked battlespace support elements, as well as by higher levels of on-board automation.
Fig. 22. Cockpits of the UH-1B (left) and UH-1Y (right, courtesy of Northrop Grumman Corporation)
Fig. 23. | HMI formats for the Lidar Obstacle Warning and Avoidance System; a) visible image, b) Planar display with detected obstacles along with ground profile represented as blue lines; c) enhanced format combining Lidar and FLIR imagery; d) Synthetic display format with wire obstacles.
ig. 24. | Key programs investigating MUMT technologies and concepts, based on the review from [92].
Fig. 26. | The Heterogeneous Airborne Reconnaissance Team (HART) system, courtesy of Northrop Grumman Corporation A GC Station (GCS) is typically a deployable structure (but can be either mobile or fixed) comprising multiple integrated GCU (i.e., multiple displays, controls and computer processors). An example of a GCS is the Heterogeneous Airbome Reconnaissance Team (HART) GCS (Fig. 26), which supports ground forces with near real-time intelligence, surveillance and reconnaissance information from multiple manned and unmanned platforms. The HART GCS communicates with a command and control center, which coordinates, prioritises and allocates re- tasking requests from multiple GCS. While a single remote pilot can potentially operate a GCS, the GCS would typically comprise multiple human operators, with each fulfilling a specific role. Peschel [100] has identified three generic roles - mission management and planning, platform control, as well as data exploitation, communications, or sensor operation. Similar to units, stations are typically associated with the control of a single RPA platform, but the design of GCS are also evolving to support the coordination of multiple platforms by multiple operators [101]. Due to the number of human operators available, GCS typically allow for more complex missions to be carried out compared to GCU. Additionally, while GCU are associated with control and coordinating RPA platforms via VLOS, GCS are associated with Beyond VLOS (BVLOS) control and coordination.
Fig. 25. | Handheld control unit. control of RPA platforms within the Visual Line of Sight (VLOS).
Fig. 27. NASA Global Hawk Operations Center, image courtesy of NASA. large number of computers, displays and controls. Centres integrate elements of GCU and GCS within a centralised facility, allowing a large number of human operators to carry out complex missions requiring the coordination of multiple RPA platforms. GCC are associated with the need for processing significant amounts of data in real-time, high level strategic mission planning, as well as coordinating efforts across multiple entities (such as ATM as well as other centres, stations and units).
ig. 29. Omni Band display concept, adapted from [110]. yellow bands comprise dashed and solid portions, which represent different thresholds on the minimum separation distance. The green portion of the Omni Band indicates the safe headings, which if flown, would either prevent or resolve a loss of well-clear. Not depicted in the figure are red bands, which represent the most serious case of loss of well-clear. Fig. 30 depicts another possible SAA HMI concept based on conflict probes [111]. The conflict probes are depicted as yellow and red contours in both the TSD and VSD, representing potential areas where a future loss of separation might occur. The size, shape and position of the probe contours vary with the intruder’ s distance, the rate of closure (dependent on the relative speed between the host and intruder aircraft), as well as the time to the closest point of approach. The probe contours indicate either violation (yellow) or collision (red) hazard criteria. Solid contours indicate regions lying on the host platform’s current flight path, while transparent probes indicate regions which do not. Similar to the Omni Band concept, information is presented on the CDTI- type display; colour banding (red and yellow) is used to indicate no-fly headings and altitudes. General Atomics Aeronautical Systems, with support from Delft University have been utilising the conflict probe display as part of ongoing research and
Fig. 28. Representative GC display with reconfigurable MFD, image courtesy of Presagis (http://www.presagis.com/solutions/ RPAS ground _control_station/).
Fig. 30. Conflict probe display concept, with permissions from the authors [113].
Fig. 31. Contingency route generation using the A viate flight planning toolbox. © 2016 IEEE. Reprinted, with permission, from [114].
‘ig. 32. | RPA mission planning software - UgCS, developed by SPH Engineering (www.ugcs.com) [115].
Fig. 33. | Anexample of a third-person display for multi-platform management. Image by S-PLANE Automation: www.s-plane.com.
y. 34. Illustration from Vigilant Spirit Control Station, from [117] (image used with permission from Springer). are used to avoid pop-ups and overlays that could obscure important information, as compared to “heads-up” concepts (e.g., Fig. 33 and Fig. 36) that seek to avoid moving the operator’s gaze away from the area of interest.
Fig. 35. Multi Aircraft Manager Tasking and Timeline Display, image used with permission of the author [119]. The HMI in Fig. 35 contains different windows (called vantages) supporting multi RPA control. The task and status vantage summarizes important aircraft status and task information. The workload management vantage provides a summarized timeline display of upcoming tasks. An expanded timeline vantage provides additional detail on aircraft sortie events. The quick views vantage offers a filtered view of upcoming tasks according to the task type.
‘ig. 36. | Glyph portrayal of RPA information, from [96]. Copyright © 2017 by John Wiley Sons, Inc. Reprinted with permission from John Wiley & Sons, Inc. Glyphs have also been proposed for enhancing operator situational awareness through representing information in a more compact manner. Glyphs are avatar-like icons, with dynamically- changing visual attributes to provide operators with platform or mission-specific information. For example, the glyph depicted in Fig. 36 contains six attributes: mission type, platform automation mode, mission status, fuel status, payload information and call sign. As glyphs depict information in a simplified graphical manner, the need for explicit textual information is reduced, therefore reducing unnecessary clutter. When depicted on a tactical display, glyphs can be co-located with the platform position, reducing the need for operators to scan between different displays in order to obtain necessary information. Additionally, dynamic changes in the visual attributes of glyphs provide salient cues of system state changes and allow adaptive alerting functions to be integrated within the glyph portrayal.
fig. 37. Interactions between the different human performance assessment techniques. These performance assessment techniques have been traditionally used to assess and evaluate human factors in new operational concepts, as well as to evaluate the design of systems and their associated HMI’. The interactions between the four techniques are illustrated in Fig. 37. Analytical models are usually employed in modelling and simulation efforts, typically used in the early stages of an engineering design process to allow refinement of the initial design before actual human-in-the-loop (HITL) studies. During actual HITL studies, task-based and psycho-physiological measures are respectively collected from the evaluated system and the human user. Finally, subjective assessments are collected from the human user at the end of the HITL study through surveys, questionnaires or interviews. Task- based, psycho-physiological and subjective measures are used to evaluate human performance in the engineering design as well as to improve the analytical model, thereby completing the design loop. Some prominent reviews on the use of human performance assessment techniques and metrics can be found in [171-175].
Fig. 38. | EEG systems: (a) B-Alert X24 EEG System, Advanced Brain Monitoring, Inc, Carlsbad, California, USA; (b) Brain Products ActiCap X press; (c) DSI 10/20, Quantum A pplied Science and Research (QUASAR) Inc., San Diego, California, USA. fNIRS system: (d) PortaLite, Artinis Medical Systems, Elst, The Netherlands. TCDS system: (e) Welder TCD headband (Image credit: David A. Washburn, Georgia State University) [312]. The increasing commercial availability of mobile/wearable brain sensing devices (Fig. 38) has opened up many avenues for neuroergonomic research. Mehta [309] provides a review of nine neuroergonomic techniques for the evaluation of cognitive and physical states. Out of the nine, four were identified as the most promising due to their high portability, relatively low cost and low intrusiveness: Event-Related Potentials (ERP), Electroencephalography (EEG) spectral analysis, functional Near- Infrared Spectroscopy (fNIRS) and Transcranial Doppler Sonography (TCDS). An additional fifth technique discussed in this section is the Steady State Visual Evoked Potential (SSV EP) [310, 311].
‘ig. 39. | Comparison of the two categories of brain activity measurement techniques.
diameter), the sampling frequency of the eye and field (for head- mounted systems) cameras, the head-box region (for remote systems), monocular/ binocular eye tracking capabilities, data processing latency as well as software compatibility. Important properties for eye tracking systems include sampling frequency, accuracy, precision and latency [421]. Systems with higher sampling frequencies are able to detect more rapid eye movements (e.g., saccades and micro-saccades). In reading, saccades occur with a latency of around 30 ms to 40 ms, so systems with sampling frequencies of 50 Hz to 60 Hz will only register the motion with one to two samples, while systems with sampling frequencies below 30 Hz are unable to reliably detect these movements [421]. Micro-saccades have the same dynamic characteristics of saccades but at a factor of 50 times smaller [422]. However, the durations of micro-saccades are only somewhat smaller than saccades. In practice, sampling frequencies of no lower than 200 Hz are used in micro-saccade research [421].
The review of adaptive HMI’ concepts presented in Section 4 allows some general observations on the characteristics of adaptive HMI’ systems. Figure 41 illustrates the general layout of an adaptive HMI’ system. The system comprises three basic, interacting modules. The human performance module assesses the human operator and, using specific human performance models, provides estimates of human performance and/or cognitive states. The task management module monitors the external environment and on-board system parameters, prioritising and tracking the progress of current tasks, as well as planning and scheduling future tasks. The adaptation engine receives inputs from the human performance and task management modules and reconfigures the HMI’ according to a set of decision logics.
‘ig.44. | Typical process flow of a task management module.
Fig.42. A classification-based approach for EEG-based workload prediction, adapted from [439]. approaches to accommodate an arbitrary number of input observables and output states, providing greater flexibility if the model’s complexity needs to be scaled up. The system is able to adapt and learn from empirical data through automatic tuning of classification parameters, allowing changes within individuals (e.g., learning or ageing effects) or differences between individuals (e.g., in personality or abilities) to be captured. However, most classification-based systems function as black boxes, providing output that is difficult to interpret and diagnose. The use of statistical or probabilistic methods lacks theoretical support from underlying cognitive principles. Additionally, classification-based systems require initial calibration on existing empirical data, making it difficult to set-up without prior data.
ig.43. Block diagram of SOAR 9, adapted from [446]. Model-based approaches place greater emphasis on modelling the processes underlying human performance than on identifying patterns from collected data. Specific aspects of human cognition such as long-term memory, working memory and perception are modelled. Model-based approaches have been used to assess cognitive workload, attention and awareness, but are also useful in evaluating higher-level processes such as planning, decision making or agent-interactions. Cognitive architectures (Section 5.3) such as ACT-R, MIDAS and SOAR (Fig. 43) are feasible candidates for a system implementation as a human performance module - past research has already explored the use of cognitive architectures in the design of autonomous systems [176, 302].
‘ig. 45. State diagram representation of human-machine behaviour, adapted from [448].
1. Introduction automated systems, which promotes more effective human- machine teaming.
‘able 4. Different concept of operations in RPAS operations and their associated HMI’ concepts
Table5. Alert categories, from [120]. Alerting systems can present a combination of visual, audial, or haptic information to the operator. Warnings and cautions usually include a combination of visual and audial elements. A consistent alerting philosophy should be present in the design of all alerts - consistency should be enforced for the prioritisation, inhibition and recall of alerts as well as the alert presentation format. Alerts should also be designed to avoid false and nuisance alerts while providing reliable and accurate alerts to the flight crew. Alerts should be managed with a prioritisation scheme. The prioritisation scheme should take into account the urgency of flight crew awareness and response as well as the alert modality. The prioritisation scheme should show that any delayed and inhibited alerts do not adversely impact safety. If multiple alerts are presented at the same time, they should be clearly distinguishable and understandable to the flight crew, as well as presented in a hierarchy to show the common root fault.
fable 8. | Cardio-respiratory measures for assessing human cognitive performance and expected response to increased task load. Measures from the cardio-respiratory system mainly comprise heart and respiratory activities. Cardio-respiratory measurement techniques are less intrusive than the techniques used to measure brain activity. Heart activity is typically measured with an electrocardiogram (ECG), while respiratory activity is measured with strain gauges. Advances in remote imaging systems provide the potential for cardiorespiratory activity to be monitored remotely [389-391]. Due to the relative ease of data collection
Hye activity has emerged as a popular measure for assessing human cognitive performance (Fig. 40) (i.e., eye movement and pupillometry). Gaze-control technology was reviewed by in a recent FAA report [408]. The report found that gaze-control technologies were sufficiently mature to be incorporated into flight deck HMI’. A number of applications were identified, ranging from eye-based input, to activity recognition, to post- flight diagnostics. However, a number of research challenges were identified, including the chance of inadvertent function activation, the accuracy and repeatability of eye measures, as well as pilot expertise affecting eye activity. Peysakhovich suggested that the integration of eye-tracking technology could be phased out over four stages with increasing on-board autonomy, namely pilot training, on-board gaze recording, gaze- based flight deck adaptation and finally gaze-based aircraft adaptation [409]. Table 9 provides a summary of the parameters used to assess eye activity, based on a review of the literature [410-415]. Specific metrics such as dispersion [416], the explore/exploit ratio [417], gaze entropy [410], Nearest Neighbour Index [418] and Index of Cognitive Activity [419] are also included in the table with reference to their respective sources. Another significant review was performed by Glaholt [413], which included 78 aviation-based empirical studies. Merchant and Schnell [420] identify a number (27) of eye tracking equipment and vendors. Table 10 presents an updated list featuring some additional vendors and systems. The table describes the tracking accuracy (pupil, gaze, head or pupil
Table 10. List of eye-tracking systems and their relevant specifications.
Table 11. HMI’ adaptation strategies, adapted from [432] Systems Design for Adaptive HMI2
Table 12. The Level of Automation Taxonomy (LOAT), from [457].
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International Journal of Engineering & Technical Research (Vol. 2 Issue 4) April -2014
by International Journals for Researchers [ER Publication, WOAR Journals, IJEAS and IJEART]

2014, Engineering Research Publication

Power quality has become an important factor in power systems, for consumer and household appliances with proliferation of various electric and electronic equipment and computer systems. The main causes of a poor power quality are... more
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MANAGING AND EVALUATING THE IMPACTS OF LARGE RESEARCH PROGRAMMES
by Trevor Wren BSc MA MPhil PhD MBA

2009

This research was funded by, and implemented within, a UK national research organisation, while the author was the KE and Impact Evaluation Manager responsible for assessing impact on a £15m UK government-funded research programme. This... more
earlier and later stages, linking idea generation to the marketplace:
TABLE 3.1: Grounding methodology and methods in literature best practice
Table 3.2: Three proposed methods within the methodology
TABLE 3.3: Analysis summary for each proposed method
2D analysis matrix with possible assessment ‘grades’ is exemplified in the table below.
TABLE 8.1: Complimentary aspects of the different methods
TABLE 10.1: Method improvements suggested after trial
data matrix analysis, gives the table below gives some case examples. information needed, and enable easier communication of the programme as a whole.
impact research and assessment may differ by impact mechanism (to be developed).
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IJETR2457.pdf
by International Journals for Researchers [ER Publication, WOAR Journals, IJEAS and IJEART]
— Theory of Eight Palaces (八宫) or Eight Veins (八 脉) is useful in understanding economic disease. By using mathematical reasoning based on Yin Yang Wu Xing Theory in Traditional Chinese Economics (TCE), this paper demonstrates the... more
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Photocopiable Glossary of Aviation Terms Word Definition and examples Resolution
by Toni Lazarovski
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International Journal of Engineering & Technical Research (Vol. 2 Issue 8) August -2014
by International Journals for Researchers [ER Publication, WOAR Journals, IJEAS and IJEART]

2014, Engineering Research Publication

The overall system block diagram is shown in figure.1. In the first part the original diabetic database consisting of the entire 8 feature is shown. This dataset will be utilized for all the classification tasks throughout the study. The entire experimental work can be divided into two major sections:
Figure 2. PCA with REP, Matlab GUI
tis ae a We FASOOUGLTIL TLVMICSOUL FTaWakt HuUllial ivitoila, Pawan Kumar Mishra is a Graduate in Computer Science Engineering from Dr. B.R Ambedkar University, Agra, ana Post graduate in Computer Science Engineering from, Uttarakhand Technical University, Dehradun. Presently he is pursuing PhD in Computer Science Engineering from, Uttarakhand Technical University, Dehradun. His area of interest include Data Structure, Object Oriented System, D.B.M.S, Mobile Computing, Software Engineering. TORRE Mir. larun jhaidiyal, larun jnaidlyal 1s a Graduate in Computer Science Engineering from Uttarakhand Technical University, Dehradun. Presently he is pursuing M.Tech. in Computer Science Engineering from Uttarakhand Technical University, Dehradun. His area of interest include Autamata, Compiler, Computer Networks, Data Structure & Data Mining. REFERENCES
From the simulation results it clear that PCA with SVM approach is better than PCA with REP approach. Table 1- Comparision Of Different Classifier Techniques.
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The Quest for Safety Amid Crowded Skies
by gines miguel franco torregrosa
As this sampling-far from a complete range-of NASA work in aeronautics indicates, the Agency and its aeronautics staff spread across the Nation maintain a lively interest in the future of flight, benefitting NASA's reputation earned in... more
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Aerodynamic technologies to improve aircraft performance
by eusebio valero

2013, Aerospace Science and Technology

An Air Transport System has become an indispensable part of Europe's economic infrastructure. The Commercial Aeronautics Sector is well aware that it has to find an acceptable balance between the constant fierce competitive pressures upon... more
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AS5210: AERODYNAMIC DESIGN DESIGN OF RC AIRCRAFT
by Saurabh Deshmukh
The required mission profile for the RC model airplane to be designed is given in Figure 1.1 below. 1.3MISSION PROFILE
Figure 8.4 Subsonic maximum lift of high aspect ratio wings [11
From [1] for NACA 4 digit airfoil AY= 26 t/c. For the chosen airfoil NACA 2310, AY= 26 x 0.1 =2.6
From Figure 8.6 for AY= 2.6 and A; -=0, Aa, =0.8' The angle of attach for maximum lift is defined as: 8.2.4 Calculation of a&cimax
Table 7.1. Historical Data for propeller selection of RC aircrafts B)
Table 5.3 Tail Volume Coefficient ©! Lyr= 0.6 x 1.174 = 0.7044 m
5.3. SUMMARY OF MAJOR DIMENSIONS OF THE AIRCRAFT Span of the horizontal tail = Sy7/Cyr
Figure 2.1. Structural or empty weight fraction plot Using Table 2.1, following two graphs, comparing the structural (also called empty weight) weight fraction and the powerplant weight fraction to the flying weight or AUW (i.e. All-up weight termed as AUW, since for electric powered RC airplanes weight during the flight remains same), respectively have been plotted.
Figure 3.1, shows the profile of the selected airfoil designated as NACA 2310. The aerodynamic and geometric characteristics of NACA 2310 airfoil are as follows:
Figure 3.4: Drag Polar For NACA 2310 In The Operating Envelope Figure 3.3: Lift Curve For NACA 2310 In The Operating Envelope
Figure 3.5: Aspect Ratio - Historical Data Plot
Figure 3.6: Wing Loading - Historical Data Plot
surveillance. The camera best suited for the mission’s requirements is the Tau 320. The payload for the aircraft is chosen as an infra-red camera to serve the purpose of forest arez 3.4.1 Payload Weight Calculation
Figure 3.8 C; v/s Cp for different velocities
Using Table 3.5 C, v/s Cpand Power v/s Velocity can plotted:
Figure 3.10 13 X 8 NEW APC Thin Electric Propeller
Figure 3.11. Silver Series 45A Brushless ESC (d) ELECTRONIC SPEED CONTROLLER (ESC)
Figure 3.9 Electrifly 4S 2200mAh LiPo (b) BATTERY — Electrifly 4S 2200mAh LiPo
Figure 4.1The Landing Path and Landing Distance[1] Assuming, flare velocity, Vp = 1.23Vetai?and load factor n=1.2"!, we get
We know that 4.3 THRUST-TO-WEIGHT RATIO CALCULATION We assume h,p= 1 m as chosen for landing. Thus,
We first convert the (T/W) for climb conditions to the (T/W) for Vmax. The power remains constant 4.5.2. Wing Loading calculation for Climb On re-arranging we get a quadratic in (W/S)
® Low Density = 75 kg/m (0.0027 Ib/in?); Medium Density = 150 kg/m? (0.0054 Ib/in?); High Density = 225 kg/m3 (0.0081 Ib/in3) Balsa wood Properties Guide Balsa wood is the softest commercial hardwood. Balsa wood's distinct properties are directly related to a unique cell structure in the material. Because of the amount of empty space within the cells of the tree, those cells are able to compress and stretch without binding or breaking. Despite the low weight and soft texture of balsa, it is one of the strongest wood types to be found. The cell's holes that are left by the drying process have a large surface that builds a sturdy support system for the exterior of the wood ™,
Lift generated by one wing = L/2 =Wo, Figure 5.1 Load Distribution with respect to Planform Shape
From the above data we get the plot: Figure 5.3 Plot of fuselage length v/s weight
5.2.3.1 Vertical Tail Sizing the wing span by. This leads to the vertical tail volume coefficient defined as:
All the required dimensions for the aircraft have been shown in Figures 6.2 and 6.3 6.3 WETTED AREA ESTIMATION For wing and empennage surfaces, the wetted area is related to the plan-form area. It is a fraction
Figure 7.1 Cross-section of a propeller[1] 7.1 INTRODUCTION The primary purpose of the propeller is to convert the power from the battery and motor power slant to axial thrust through torque transfer to the propeller. Propellers may be classified as to whether the blade pitch is fixed or not. The demands on the propeller differ according to circumstances. For example, in take-offs and climbs more power is needed, and this can best be orovided by low pitch. For speed at cruising altitude, high pitch will do the best job. Propellers are orimarily classified as: CHAPTER 7: PROPELLER SELECTION
Figure 7.5 Special thin, light, and wide blade (12 *%6 APC Slowflyer prop)[3 3. Propeller pitch (P) = 6 inches 7.4 PERFORMANCE RELATIONSHIPS
The standard variation of propeller parameters is given below in Figures 7.6 and 7.7
Figure 7.7. Variation of J with C, at different P/D ratios 5] Figure 7.6. Variation of | with Cr at different P/D ratios 51
The uncambered wing has no lift at zero angle of attack. Maximum lift is obtained at the stall angle of attack, beyond which the lift rapidly decreases. When a wing is stalled, most of the flow over the top has separated. The slope of the lift curve is essentially linear except near the stall angle, allowing the lift coefficient below stall to be calculated simply as the lift curve slope times the angle of attack. At the stall, the lift curve becomes non-linear. The effect of Mach number on the lift- curve slope is shown in Figure 8.1. The maximum Mach number for maximum velocity of the aircraft is 0.083. shown in Figure 8.1. The maximum Mach number for maximum velocity of the aircraft is 0.083. Figure 8.1 Lift curve slope v/s Mach number
Figure 8.2 Taper ratio correction for low aspect ratio wings Where C, is a constant based on the taper ratio and A; is the angle at the leading edge. For C,;=0 and A; -=0 A < 3. For our aircraft A= 5.35 > 3.
Figure 8.3 Airfoil leading edge sharpness parameter I! = 1.467 for the airfoil chosen. mi Here C max is the airfoil maximum lift coefficient at M=0.2
that for low subsonic speeds there is an asymptotic lower limit to the parasite drag. Therefore, zero-lift drag coefficient, Cp9= 0.01759.
Figure 8.8 shows the new drag polar.
We first convert the (T/W) for climb conditions to the (T/W) for Vmax. The power remains constant.
Figure9.2. Top view of the aircraft in cartesian system
A=0.409
Thus, Centre of Gravity from fuselage nose tip Zcq = D = 125.94 mm
The landing gear locations from CG (/,, for main gear and /,, for nose gear),base (b) and track (t) are defined as shown below;
Figure 10.1. Position of %4 root chord I
Figure10.2: Theoretical lift increment Figure10.3:Empirical correction for for plain flaps plain lift increme!
10.13 CONCLUSION
‘igure 11.1 Cp values for Landing Gear (Fixed Type) |!
Figure 11.3 Parasite Drag vs. Mach number The parasite drag calculated using the component build-up method is 0.02320. Thus the modified
12.11. SUMMARY AND SELECTION OF THRUST TO WEIGHT RATIO Thrust for take-off calculated in section 12.7.1= 12.409 N 12.10.4 Take-Off Table 12.2 Thrust-to-weight ratio of different segments
The curve between O and A represents the aerodynamic limit on load factor imposed by Cmax. The region above the curve OA is stall region. Horizontal line AB represents the limit load factor. On the right hand side of diagram, line BC represents the maximum diving speed. The bottom part of the diagram, given by curve OD and horizontal line DC corresponds to negative absolute angle of attack, that is negative lift and hence the negative load factor. The curve between O and A represents the aerodynamic limit on load factor imposed by Cmax. The
The right side of the V-n diagram describes about the high speed limits of the aircraft. Undesirable phenomenon such as wind gust, aileron reversal etc. will lead to structural failure of the aircraft. The speed limit is given by the red line. Under no circumstances, should this velocity be exceeded. During diving, the weight and the thrust vector are along the same direction and maximum thrust and acceleration is obtained. However, this is undesirable for smaller aircrafts and the maximum velocity is constrained by the power available.
LIST OF TABLES
Table 1.1 Design Specifications
Table 1.3 Preliminary design considerations The preliminary design considerations arrived at, after literature survey, are enlisted in Table 1.3
Table 2.1 Weight specification for RC airplanes
Table 2.2. Result table of iterative process 2.4 CONCLUSION was allowed to continue till an error of less than 0.5% had been achieved.
Table 3.1 Airfoil Data rigorous approach of custom-designed airfoil based on computational analysis has been avoided. Following Table 3.1 shows a compilation of some of NACA 4-digit family of the airfoils considered for this work, as the RC model airplane has to operate at very low speed in surveillance operation. This set of airfoils has been selected based on the high stall angle, high maximum section lift coefficient, availability of thickness for structure and ease of manufacturability. At this stage of design a more rigorous approach of custom-designed airfoil based on computational analysis has been avoided.
Table 3.2 Characteristics of NACA 2310
Table 3.5 Parameter values at different velocities Table 3.5 below tabulates the values of C,, Cp and power for different velocities.
Table 4.2 Wing Loading at different flight conditions 4.2 WING LOADING SELECTION
Table 4.4 Thrust-to-weight at different flight conditions Tabulating the values of (T/W) of other segments using this value of power The maximum power for which the aircraft has to be designed is 87.20 W which is during climb.
For a propeller-driven airplane, power to weight ratio is more relevant than the thrust-to-weight ratio /! Table 4.3 Engine power at different flight conditions 4.4 SUMMARY OF THRUST-TO-WEIGHT AND SELECTION
Table 5.1 Historic data for fuselage length and total weight of the aircraft
5.1.7 Third Weight Estimate
Table 8.1 RC Model Airplane (Wo = 1.98 kg) Component Geometric Data 8.3.6 Calculation and Summary Table of Parasite Drag The parasite drag is thus calculates as:
Table 8.2 Parasite Drag Calculation for design cruise speed of 20 m/s Table 8.3 below shows the tabulated values of Cpg against the airplane speed and the corresponding Mach number.
Table 8.5 Engine power at different flight conditions 8.9 SUMMARY OF THRUST-TO-WEIGHT AND SELECTION
Table 8.6 Summary of power requirements for different segments 8.11 SUMMARY OF POWER REQUIREMENTS
Table 9.1 CG location and weights of different components 9.2.9 Summary of the CG location of different components
Table 11.1 RC Model Airplane (Wo = 1.98 kg) Component Geometric Data Table 11.1 summarizes the geometric data of airplane components used for the estimation of drag values. Wing and empennage airfoils have t/c of 0.1.
Table 11.2 Parasite Drag Calculation for maximum cruise speed of 20m/s
Table 12.1 Wing loading for different segments Thus, the wing loading selected is the minimum from the table above Table 12.1 below summarizes the values of wing loading calculated in the above sections
Prax = 133.61 W
Table 12.3 Parameters required for V-n diagram 12.12.4 Limit load factor calculation Positive limit of the limit load Factor is given by"!
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AIRCRAFT FLIGHT CONTROL SYSTEM
by RAJENDRA PATEL
A control system is a collection of mechanical and electronic equipment that allows an aircraft to be flown with exceptional precision and reliability. A control system consists of cockpit controls, sensors, actuators (hydraulic,... more
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International Journal of Engineering & Technical Research (Vol. 2 Issue 9) September -2014
by International Journals for Researchers [ER Publication, WOAR Journals, IJEAS and IJEART]

2014, Engineering Research Publication

www.erpublication.org  Abstract-We present a practical scheme for measurement-device-independent polarization shift keying using two state polarization encoding. Most of the previous work on optical free space laser communications... more
Since the information is embedded in polarization of the photons, it is the snake and ballistic photons which are relevant to us. Diffused photons therefore are a part of noise and if more and more of the incident photons fall into the diffused regime, information gets scrambled and will be completely lost. Fig. 1. Schematic showing ballistic, diffusive and snake photons from a multiple scatterer media.
S ranges from —1 to +1 fora purely plane polarized light. The range is less than +1 for partially polarized light and the The receiver consists of another polarizing beam splitter and two Single Photon Counting Modules (SPCM), based on Avalanche Photodiodes (SensL PCD mini 0200 with sensor area 20 ym, cooled to -20°C using Peltier modules). They are labelled APD1 and APD2. The TTL pulses from these units are collected using a DAQ card (NI-PCI-6320, from National
The initial ‘information’ at the transmitter is just vertical or horizontal polarization, and hence message bits 0 and 1 correspond to State of polarization —1 and +1. At the receiver,
Fig. 3. Histograms for photon counts for calibration data. (a) and (b) are respectively APD1 and APD2 counts when only L1 is used with vertical polarized pulses. While APD 1 alone should have registered counts, APD2 also exhibits some counts, though averaged at about 15 as against average of 2000 counts for APD1. Similarly (c) and (d) are for when only L2 is used (horizontal polarization). This background counts (termed ‘leakage’) are due to dark counts, stray light as well as some imperfections in polarizing beam splitters. Since histograms do not overlap, signal can be extracted inspite of these imperfections.
Fig. 4. (a) Distribution of State of Polarization with no atmospheric effects in between. The distribution consists of two non-overlapping peaks centered at —1 and +1 respectively. (b) With smoke at OD=-25 dB in path. The range is reduced from +1. However, bit assignment will not have errors until the distributions start overlapping
We then transmitted the random sequence of 0’s and 1’s for about 50000 data bits, SOP was computed for each bit using formula 4. A histogram of different SOP values were obtained and shown in figure 4(a). It shows two peaks, centered at +1 respectively. It also shows that the distributions are not overlapping. In addition, these two peaks were fitted independently to Gaussian functions, in order to obtain the width of these distributions (figure 5). The width is a measure of polarization impurity, since pure polarized light should give a very narrow distribution. The impurity is either due to actual degradation of polarization or could also be due to measurement errors and dark counts of the APD’s. A very badly degraded polarization, or bad measurement would give
Fig. 6. Q factor for vertical polarized (left) and horizontal polarized (right) lights, in presence of smoke, as a functlion of Optical Density. Thick line is a curve fitting for a stretched exponential function, with stretch factor # =0.49. Sa NOE Bee ee Figure (6) shows Q factor for vertical and horizontal polarized lights in presence of smoke. Optical density as defined by equation (7) is used for quantifying the amount of smoke in the chamber. It can be noticed that as Optical density increases from 0 to -30 dB, the quality factor Q reduces to almost zero. This is expected since the multiple scattering by the smoke particles degrades the polarization and errors increase. W e fit this to a stretched exponential curve aexp((x — b)’) +c to fit the data, where a is the amplitude, b is the shift in x axis and f is the stretch coefficient. We find that the curve fits for 6 = 0.5.
Fig. 7. BER as computed by equation 10 for vertical polarized (left) and horizontal polarized (right) lights, in presence of smoke, as a function of Optical Density
After transmission, each bit on receiver was compared with the transmitted data and correlation between them was computed. figure 8(a) shows that despite high attenuation of light and a high degradation of the initial polarization, the actual bit error are almost zero and fidelity of transmitted data is very high, as can be seen by 1 — 1 and 0 — 0 correlations. But at higher concentrations of smoke, errors do develop, as shown by figure 8(b). It can be noted that actual BER is nearly zero for a very high amount of smoke, and suddenly increases only at a very high OD, like -30 dB. This shows that despite theoretically low values of Q and high values of BER, the actual BER is very less. It can be noted that theoretical BER is computed for On/Off keying while the experimental value is for PolSK, indicating that PolSK is much more robust ig. 9. Q factor for vertical polarized (left) and horizontal polarized (right) lights, in presence of fog, as a function of Optica Density Fog is an atmospheric condition made up of tiny water droplets, which also scatter light. Attenuation due to scattering, rather than absorption as in case of smoke is the key issue in case of fog. Multiple scattering by fog droplets also degrades polarization. Fog particles are larger than smoke particles and hence scattering cross section is higher. The polarization changes due to scattering is also different from that for smoke.
Fig. 10. BER as computed by equation (10) for vertical polarized (left) and horizontal polarized (right) lights, in presence of fog as a function of Optical Density. Solid line indicates a curve fit for exponential function.
Fig. 8. (a) Correlation between transmitted bits, obtained from physically comparing each bit. The cross-correlation is almost zero and all bits are therefore transmitted with a very high fidelity. (b) Experimentally measured BER by comparing individual bits of transmitted and received data. Thick line represents a fit to exponential function of a stretched exponential function, with stretch factor f ~ 0.5, whereas BER depends on OD with a normal exponential function. This indicates that the Q factor has a slow decrease with increasing noise levels while BER has a very sharp increase beyond a threshold noise. Asearlier, SOP was computed for each bit. A histogram of the SOP distribution shows that its range is reduced from +1 to a lesser range (figure 8). However, the distributions are still not overlapping and hence it can be safely assigned to bit 0 when SOP is negative and to bit 1 when SOP is positive.
Fig. 11. Experimentally verified BER by comparing each bit. Solid line indicates a curve fit for exponential function. As incase of smoke, we also calculated BER from comparing actual transmitted and received data. Although in case of smoke, the experimental BER for PolSK was much smaller than the theoretical BER for OOK, fog data on the other hand shows much closer comparison. This is because the scattering loss for polarization are different for smoke and fog particles.
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IJETR2454.pdf
by International Journals for Researchers [ER Publication, WOAR Journals, IJEAS and IJEART]
 Abstract— Theory of Eight Palaces (八宫) or Eight Veins (八 脉) is useful in understanding economic disease. By using mathematical reasoning based on Yin Yang Wu Xing Theory in Traditional Chinese Economics (TCE), this paper demonstrates... more
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ATPL Radio Navigation © Atlantic Flight Training
by Danny YI
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International Journal of Engineering & Technical Research (Vol. 2 Issue 5) May -2014
by International Journals for Researchers [ER Publication, WOAR Journals, IJEAS and IJEART]

2014, Engineering Research Publication

Abstract-This work has been carried out for the purpose of improving productivity and effectiveness of system for the integrated management of tools within a company, by integrating planning, inspection and tool-room functions; automated... more
Figure 1. Conceptual model. The conceptual model underlying the present study is given in Figure 1.
Figure 2, ANN-nultilavered perceptron. adaptive learning algorithm and use an information processing system composed of a large number of interconnected processing elements (neurons) working in tandem. Neural networks are made of basic units arranged in layers. A unit collects information provided by other units (or by the external world) to which it is connected with weighted connections called synapses. These weights, called synaptic weight multiply (i.eamplify or attenuate) the input information. A positive weight is considered excitatory, a negative weight inhibitory. One of the most popular architectures in neural networks is the multi-layer perceptron which is illustrated in Figure 2. Learning happens through a methodology of continuously altering the weights to achieve closer and closer values to desired output. One algorithm that performs this is known as the back propogation algorithm. The back propagation algorithm is a generalization of the least mean squre algorithm . The network weights are modified to minimize the mean squred error between the desired and the actual output of the network.The network is trained using a training data set where the inputs and output Value are Known. After the tranning is completed, the weights are frozen and the models can be used be prediction or outputs for new set of inputs Values.
A typical ANFIS consists of five layers, which perform different actions in the ANFIS are detailed below. For simplicity, we have illustrated a system that has two inputs x and y and one output Z. The rule base, for illustrative purposes consists of two if-then rules of the Ta-kagi-Sugeno type.
converts the crisp inputs into “degree of match “with the linguistic values like high or low etc., and 5) a de fuzzification interface that reconverts to a crisp output.
Table 1: Sample Dataset
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DISEÑO SIMPLIFICADO Y CONSTRUCCIÓN DE UNA AERONAVE NO TRIPULADA (UAV) A ESCALA PARA COMBATIR INCENDIOS - FIREHUNTER
by Nicolás Castro Latorre
Present paper is the summit report of the simplified design and prototype construction of the FireHunter UAV for firefighting purposes as applied academic project comply of the Aeronautical Engineering Students enterprise; ColTech... more
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THEORY AND ANALYSIS OF ELASTIC PLATES AND SHELLS Second Edition
by Tareq Dahb
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IJETR2458.pdf
by International Journals for Researchers [ER Publication, WOAR Journals, IJEAS and IJEART]
 Abstract— Theory of Sixty JiaZi (60 甲子) is useful in understanding economic disease. By using mathematical reasoning based on Yin Yang Wu Xing Theory in Traditional Chinese Economics (TCE) , this paper demonstrates the treatment... more
Figure 4. Ten Heavenly Stems Hidden behind Twelve Earthly Branches
Figure 5. Relations between Eight palaces and Six-fus
The steady multilateral system
Figure 3. Relations of twelve Earthly Branches
Table 1. Sixty JiaZi naming of ten heavenly stems,twelve earthly branches and Yin Yang attribute Table 2. Sixty JiaZi meaning of ten heavenly stems of state, good & bad and comprehensive name
Table 3. Sixty JiaZi meaning of philosophy,good & bad and comprehensive name
Assume that M_ or M ; as issued in the circulation of generalized money, Gross Domestic Product (GDP) as G or G’ and the consumer price index (CPI) as C or C’' for today and last year respectively, the actual need of money in real terms in the circulation F = M3 x(G/ G') for last year's price level. Then the inflation rate of M ‘ is |( M, =f ait M 4 , the inflation rate of GDP is (G-G')/G', and the annual finance inflation rate can be measured by (M, oy E )/ a Table 4. Inflation Rates in Finance, GDP and CPI
Table 5. Inflation Rates in PPI,RPI,GBR and AAF
Let x‘ be the one of PACGGF inflation rates for any j (1 < i < 6) of the steady multilateral system. Denoted the parameters of the normal range by @ byt, 1 = 1/2j3,4,5,6; The Hexagram-image of the Eight-Palaces or Eight Veins or Eight Extra Meridians is (hye fos Sys Sas tes Sq): Wespectively. where f =sign(x'=2+@' =), 7=1,2,3,4,5,6.
Instead of the one x’ of PACGGF inflation rates to x’ of the RPI inflation rate, the Numbers similarly. Table 7. Hexagram-images of Eight Palaces (II)
Tables 8. Sixty JiaZi meaning of ten heavenly stems, branches to hide stems, synthesized attribute and integration attribute
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Global navigation satellite systems performance analysis and augmentation strategies in aviation
by Roberto Sabatini and 
1 more
Subramanian Ramasamy
In an era of significant air traffic expansion characterised by a rising congestion of the radiofrequency spectrum and a widespread introduction of Unmanned Aircraft Systems (UAS), Global Navigation Satellite Systems (GNSS) are being... more
In terms of WAAS integrity features, Message Types 27 and 28 are particularly important. Type 27 Messages may be transmitted to correct the cUDRE in selected areas. Each Type 27 Message specifies SUDRE correction factors to be applied to integrity monitoring algorithms of users when inside or outside of a set 0 geographic regions defined in that message. SUDRE indicators are associated with the SUDRE values listed in Table 17 that multiply the model standard deviation defined using the UDRE parameters in the WASS Types 2-6 and Type 24 Messages. Table 16. SBAS Operational Classes [41].
Fig. 1. GNSS segments. includes the large variety of GNSS receivers developed for air, ground and marine navigation positioning applications. The control segment includes one or more Control and Processing Stations (CPSs) connected to a number of Ground Monitoring Stations (GMSs) and antennae located around the globe for Telemetry, Tracking and Command (TT&C) signals down/uplink and navigation/integrity signals uplink to the satellites. The GMS antennae passively track all GNSS satellites in view collecting ranging signals from each satellite. This information is passed on to the CPS where the satellite ephemeris and clock parameters are estimated and predicted. Additionally, satellite integrity data are analysed and appropriate integrity flags are generated for faulty/unreliable satellites. The ephemeris/clock and integrity data are then up-linked to the satellite for retransmission in the navigation message. The satellite clock drift is corrected so that all transmitted data are synchronised with GNSS time.
Fig. 2. Navigation solution in the ECEF co-ordinate system (at time zero, the Xz axis passes through the North Pole, and the Y g axis completes the right-handed orthogonal system. satellites. The GNSS receiver’s computer may be programmed to solve directly the navigation equations in the form given above. However, the computation time required to solve them may be too long for many applications. As an alternate approach, these equations may be approximated by a set of four linear equations that the GPS receiver can solve using a much faster and simpler algorithm. The system of equations (9) can be rewritten in the form:
where ®{(r) is the phase reading (phase at receiver A of the signal from satellite i at time rt); ©*(t) is the received signal (phase of the signal as it left the satellite at time t) and @,(z) is the generated signal phase (phase of the receiver’s signal at time tz). If p4(t) is the range between receiver and satellite, we have:
Fig. 3. Error components in ephemeris estimation. Corrections to the drift of the satellite atomic clocks are computed by the CPS and then broadcasted to the users in the navigation message. The effect of satellite clock offset is negligible in most positioning applications (using the polynomial coefficients corrections computed at the CPS it is possible to reduce this error down to 1 part per 10’’). The residual error is due to the fact that corrections from the CPS are periodic and not continuous. Group Delays are the delays typical of the satellite electronic circuits. They are estimated on the ground before the satellites are launched and corrections are included in the navigation message. ionosphere can be represented by a thin shell at an elevation of 350 km, the obliquity can be approximately expressed as a simple function of the elevation angle in degrees (£) of the satellite at the receiver’s antenna [14]:
values of less than 2-3 metres are the norm and upper values of 15 metres are rarely observed. For carrier phase, the maximum theoretical multipath error is a quarter of the wavelength. This equates to about 5 centimetres for the GPS L1 and L2 frequencies, although typical values are less than 1 centimetre [27]. Multipath can be accurately modelled and removed only at static points, by taking observations at the same points and at the same hour on consecutive days. This, however, is possible in a dynamic environment. Other techniques use Signal-to-Noise Ratio (SNR) and signal phase/frequency information to detect and quantify multipath.
Fig. 5. UERE variation with time in normal and extended operations [31]. We! tip 1 as future users of GALILEO and BEIDOU, will be able to use mu e frequencies to compensate for the ionospheric errors and thereby to achieve lower UEREs. In GNSS systems the as UE ast shows variation with time [31]. In GPS normal operations, the time since up t up oad the RE values are strongly dependent on the time elapsed since the from the control segment. As an example, Fig. 5 GPS Standard Positioning Service (SPS) UERE oad satellite. is limited to no more than one day. The smallest UERE and best SIS accuracy will generally occur immediately after an upload of fresh NAV message data to a satellite, while the largest UERE and worst SIS accuracy will usually be with the stalest NAV message data just prior to the next upload to that
where an error vector has been added to account for pseudorange measurement noise plus model errors and any unmodelled effects (e.g., SA), and f;; is the direction cosine of the angle between the LOS to the i® satellite and the j” co-ordinate.
There is a proportionality between the PDOP and the reciprocal value of the volume V of a particular tetrahedron formed by the satellites and the user position (Fig. 6). Fig. 6. PDOP tetrahedron construction.
Fig. 7. “Cut and fold” tetrahedron for PDOP determination. At a particular time and location, the four satellites shown in Fig. 6, have determined elevations and azimuths with respect to the receiver. From these satellite positions the components (x,y,z) of the unit vectors to the satellites can be determined. Using the Pythagorean theorem, the distances between the ends of the unit vectors can be determined. Knowing these distances, a tetrahedron can be constructed (Fig. 7) and the four altitudes of the tetrahedron can be measured.
Fig. 8. Navigation accuracy in aviation applications. navigation and flight control systems. This second type of accuracy is called Total System Error (TSE) and is measured as deviation from the required flight trajectory. In large commercial aircraft and several military aircraft, the required flight trajectory and associated guidance information is computed by a Flight Management System (FMS) and constantly updated based on ATM and weather information. The navigation system estimates the aircraft state vector (i.e., position, velocity, attitude and associated rates), determines the deviation from the required flight trajectory and sends these information either to the cockpit displays or to an Automatic Flight Control System (AFCS). The error in the estimation of the aircraft’s position is referred to as avigation System Error (NSE), which is the difference between the aircraft’s true position and its displayed position. The difference between the desired flight path and the displayed position of the aircraft is called Flight Technical Error (FTE) and accounts for aircraft dynamics, turbulence effects, human- machine interface, etc. As illustrated in Fig. 8, the TSE is obtained as the vector sum of the NSE and the FTE.
Fig. 9. GNSS satellite obscuration analysis. It should be noted that even high performance avionics GNSS antennae have gain patterns that are typically below -3dB at about 5 degrees elevation and, as a consequence, their performance become marginal below this limit (Fig. 10).
Fig. 10. Avionics antenna gain pattern (L1 frequency).
Fig. 11. GPS satellite antenna coverage. The link budget calculated from equation (55) only refers to the direct GNSS signal received from a satellite. Multipath effects, which are due to the geometric and reflective characteristics of the environment surrounding the GNSS antenna are not included in this calculation and are discussed separately. The L-band antenna on-board a GNSS satellite is designed to radiate the composite L-band signals to the users on and near the Earth. In the case of GPS (Fig. 11), the satellite viewing angle from edge- to-edge of the Earth is about 27.7 degrees [46]. The satellite antenna is designed to illuminate the Earth’s surface with almost uniform signal strength. The path loss of the signal is a function of the distance from the antenna phase centre to the surface of the earth. The path loss is minimum when the satellite is directly overhead (90° elevation), and is maximum at the edges of the coverage area (satellite at the horizon).
Geographically, there are two zones of intense scinti high latitudes and the other cen equator as shown in Fig. 12. observed in these two sectors scintillation occurs exceptional storms. In the equatorial sector, maximum of activity. For equatorial scintillation, tred within +20° of , while in the mid y, such as during lation, one at the magnetic Severe scintillation has been dle latitudes geomagnetic there is a pronounced night-time peak activity around the vernal equinox and high activity at the autumnal equinox have been observed. Fig. 12. Depth of scintillation fading at 1.5 GHz during solar maximum and minimum years [47].
Fig. 13. Reference geometry for Doppler shift analysis. where v is the aircraft velocity vector, E,,is the elevation angle of the n™ satellite, y, is the relative bearing of the aircraft to the satellite and y,,’is the azimuth of the LOS projection in the antenna plane. From equations (67) and (68), the aircraft-satellite relative geometric conditions that maximise or minimise Doppler shift can be determined. In particular, combining the two equations, the Doppler shift is given by:
The above equation shows that both the elevation angle of the satellite and the aircraft relative bearing to the satellite affect the magnitude of the Doppler E,, lead to increases in Dopp or decrements in Doppler shift de and the direction of the aircraft Fig. 14, itis evident that a re shift. er shift, while y,,’ drives increments In particular, reductions of pending on the size of the angle velocity vector. By inspecting ative bearing of 90° and 270° would lead to a null Doppler shift as in of the aircraft velocity vector (¥q in this case) in the LOS to the satellite. However, in all other cases (i.e, 7 #7), such a component would be present and this would lead to increments or decrements in Doppler shift depending on the relative directions of the vectors ¥ and v,. This fact is better shown in Fig. 14, where a steady flight without loss of generality is assumed. this case there is no component As the time required for typical aircraft heading change manoeuvres is much shorter than the timeframe associate to significant satellite constellation changes, each satellite can be considered stationary in the body reference frame of the manoeuvring aircraft. Therefore, during manoeuvres leading to Doppler shift, an initial aircraft velocity vector 7 can be considered to define path constraints that avoid further signal degradation or loss. This can be performed by imposing that the aircraft increases the heading rates towards the directions ¥ or - vq and minimises the heading rates towards the directions ¥, and -Vy. The choice of % or -% is based simply on the minimum required heading change (ie., minimum required time to accomplish the correction). Regarding the elevation angle (E dependency of Doppler shift, equation (69) shows that minimising the elevation angle becomes an objective also in terms of Doppler trajectory optimisation.
Fig. 16. Variation of A, as function of the angle f. where A, is the direct amplitude and f is the signal amplitude, A,, is multipath signal phase of the multipath. Fig. 16 shows that both the multipath phase @ and the multipath amplitude affect the received signal. Therefore, we require a multipath model to simulate these two fac airframe and from Aeronautical Multipath tors, considering the reflections from the the ground. The commonly adopted Channel (AMC) model developed during the ESA-SDS research 50, 51].
From Fig. 15, the SNR and phase error for a single refection can be represented as a function of direct and multipath signal amplitudes and the multipath relative phase P [49]:
Fig. 17. AMC model structure. Fig. 18 shows the geometric reflection model. The incoming wave is emitted from point T, R is the receiver location and S is the reflection point. V is a defined point on the reflecting surface and n stands for a unit vector normal to the surface.
Fig. 18. Geometric reflection model. In ray-tracing the reflection point S and the defined point V should satisfy the equation:
a A a The system consists of a Reference Receiver (RR) located at a known location that has been previously surveyed, and one or more DGNSS User Receivers (URs). The RR antenna, differential correction processing system, and data link equipment (if used) are collectively called the Reference Station (RS). Both the UR and the RR data can be collected and stored for later processing, or sent to the desired location in real time via the data link. DGNSS is based on the principle that receivers in the same vicinity will simultaneously experience common errors on a particular satellite ranging signal. In general, the UR (mobile receiver) uses measurements from the RR to remove the common errors. In order to accomplish this, the UR must simultaneously use a subset or the same set of satellites as the reference station. The DGNSS positioning equations are formulated so that the common errors cancel. The common errors include signal path delays through the atmosphere, and satellite clock and ephemeris errors. For PPS users, the common satellite errors are residual system errors that are normally present in the PVT solution. For SPS users, the common satellite errors also include the intentionally added errors from SA. Errors that are unique to each receiver, such as receiver measurement noise and multipath, cannot be removed without additional recursive processing (by the reference receiver, user receiver, or both) to provide an averaged, smoothed, or filtered solution. Various DGNSS techniques are employed depending on the accuracy desired, where the data processing is to be performed, and whether real- time results are required. If real-time results are required then a data link is also required. For applications without a real-time requirement, the data can be collected and processed later. The accuracy requirements usually dictate which measurements are used and what algorithms are employed. In the case of Differential GPS (DGPS), accuracy is independent of whether SPS or PPS is being used, although real-time PPS DGPS can have a lower data rate than SPS DGPS because the rate of change of the nominal system errors is slower than the rate of change of SA. However, the user and the Reference Station must be using the same service (either PPS or SPS). The clock and frequency biases for a particular satellite will appear the same to all users since these parameters are unaffected by signal propagation or distance from the satellite. The pseudorange and delta-range (Doppler) measurements will be different for different users because they will be at different locations and have different relative velocities with respect to the satellite, but the satellite clock and frequency bias will be common error components of those measurements. The signal propagation delay is truly a common error for receivers in the same location, but as the distance between receivers’ increases, this error gradually de-correlates and becomes independent. The satellite ephemeris has errors in all three dimensions. Therefore, part of the error will appear as a common range error and part will remain a residual ephemeris
where P” is the pseudorange measurement, p? denotes the geometric distance between the stations and satellite, dt; denotes the receiver’s clock offsets, d;,, denotes the receiver’s hardware code delays, d?,, denotes the multipath of the codes, Ag, denotes the measurement noise and c is the velocity of light. Equation observab above eq Hence, differenc differenc differen Double receiver (104) represents the single difference pseudorange e between receivers. There are four unknowns in the uation assuming that the co-ordinates of station k are known and that the difference in clock drifts is one unknown. four satellites are required to provide four single e equations in order to solve for the unknowns. Single es with code observations are frequently used in relative tial) navigation. Using all pseudoranges shown in Fig. 22, differences are formed between receivers and satellites. differences are constructed by taking two between- single differences and differencing these between two satellites. This procedure removes all satellite dependent, receiver dependent and most of the atmospheric errors (if the distance between ia the two receivers is not too large). The derived equation
Fig. 21. GNSS integration architectures: a. OLGS, b. CLGS and c. FIGS [107]. The other classification commonly in practise to define the integration techniques is given by:
As mentioned before, step anomalies will normally cause carrier- phase DGNSS receivers to lose lock on the carrier phase, causing the reference and user receivers to reinitialise. UR noise, processing anomalies, and multipath at the user GPS antenna cannot be corrected by a DGNSS system. These errors are included in the overall DGNSS error budget. Errors in determining or transmitting the satellite corrections may be In the aviation context, various strategies have been developed for increasing the levels of integrity of DGNSS based navigation/landing systems. These include both Space Based Augmentation Systems (SBAS) and Ground Based Augmentation Systems (GBAS) to assist aircraft operations in all flight phases (Fig. 22). Fig. 22. GBAS and SBAS services. Adapted from [111].
Fig. 23. FDE Events [36].
Fig. 24. Protection levels and alert limits.
Fig. 25. Protection levels and alert limits. The circumferences shown in Fig. 25 have their centres at the estimated aircraft position and their radii are the system PLs green) and the operation ALs (red). The aircraft shape represents the actual position, so the navigation system error is the distance from the centre of the circumferences to this shape. According to the previous definitions, situations B, C and F represent integrity events. In particular, case B is a MI event and cases C/F are HMI events. The desirable situation is represented by case A. Particular attention should be given to situations B and especially C, which is the most feared situation, as the system does not alert the user for the unavailability of the system and, depending on the on-going flight operational task; this may lead to a SoL risk.
Fig. 27. WAAS Architecture and Operational Environment [63].
lonospnenc delaY 15S One OL We Major efor sources OF pseudorange. The WAAS ionospheric delay corrections are broadcast as vertical delay estimates at specified ionospheric grid points (IGPs), applicable to a signal on L1. The WAAS Reference Stations (WRSs) measure the slant ionospheric delays to all satellites in view. These measurements must be translated into a form that can be applied by the user because the user will have a different line of sight to the satellite than the WRSs. WAAS uses a two-dimensional grid model to represent the vertical ionospheric delay distribution [47]. Ten different bands are used to allow a regular spacing of IGPs. The 1808 possible IGP ocations in bands 0-8 are illustrated in Fig. 28 (there are 384 additional IGP locations in bands 9-10 not shown). In the bands 0-8, the IGPs are spaced 5° apart from each other both in ongitude and latitude for latitudes up to N55° and $55°. Above 59° and up to 75° in latitude the IGPs are spaced 10° from each other both in longitude and in latitude. At S85° and N85° (around the poles), the IGPs are spaced 90°. In the bands 9-10, the IGP grid at 60° has 5° longitudinal spacing, increasing to 10° spacing at 65°, 70° and 75°, and finally becoming 30° at N85° and S85° round the poles. To accommodate an even distribution of bands, IGPs at 85° are offset by 40° in the 0-8 bands and 10° in the 9-10 bands. The IGPs error data are transmitted in Message 26 and denoted in terms of GIVEI (Grid Ionospheric Vertical Error Indicator), which corresponds to specific values of Grid Ionospheric Vertical Error (GIVE) and GIVE variance (07; giz). The GIVEIi, GIVEi and the 0”; ¢;y__ are listed in Table 15. Fig. 28. WAAS IGPs for bands 0-8 [41].
where 0; ionogria iS the model variance of inospheric vertical delays at an IGP. As indicated in the WAAS MOPS, a dedicated model can be used to calculate both fast and long-term correction degradation components (for inclusion in Message Types 7 and 10).
and the observation matrix G consists of N rows of LOS vectors from each satellite augmented by a “1” for the clock. Thus, the ith row corresponds to the ith satellites in view and can be written in terms of the azimuth angle Az; and elevation angle 6;. The i” row of the G matrix is: where G is the observation matrix, W is the weighted matrix, Seasti 1S the partial derivative of position error in the east irection with respect to the pseudorange error on the ith satellite, a a a d Snortni 1S the partial irection with respect irection with respect and sy; is the partial pseudorange error on the i" satellite. The weighted matrix W is efined as: derivative of position error in the north to the pseudorange error on the i" satellite, Sy, is the partial derivative of position error in the vertical to the pseudorange error on the i" satellite derivative of time error with respect to
have been already developed by the ICAO Navigation System Panel (NSP) but not yet included in Annex 10, waiting for further developments to take place in the aviation industry. As shown in Fig. 31, GBAS utilises three segments: satellites constellation, ground stations and aircraft receiver. The GBAS ground station is formed by reference receivers with their antennas installed in precisely surveyed locations. The information generated in the receiver is sent to a processor that computes the corrections for each navigation satellite in view and broadcasts these differential corrections, besides integrity parameters and precision approach pathpoints data, using a VHF Data Broacast (VDB). The information broadcast is received by aircraft in VHF coverage that also receive information from the navigation satellites. Then, it uses the differential corrections on the information received directly from the navigation satellites to calculate the precise position. The precise position is used, along with pathpoints data, to supply deviation signals to drive appropriate aircraft systems supporting precision approach operations. Although the operating principle, Signal-in-Space and performance characteristics of GBAS are completely different from ILS, the GBAS approach guidance inidications provided to the pilot are similar to the localizer and glide path indications provided by an ILS. However, compared to ILS precision approach systems, GBAS presents several distinct benefits. These include:
Fig. 32. Minimum category I, II and APV service volume (adapted from [65]).
Fig. 33. ABIA system architecture.
Targeting an ABIA implementation, a dedicated analysis is required in order to determine the flight envelope limitations associated with the use of GNSS. In particular, the following models have to be introduced [53, 54]:
Current SBAS and GBAS technologies are not able to meet the stringent GNSS data integrity requirements imposed by airworthiness regulations in some of the most demanding operational tasks (e.g. UAS sense-and-avoid). GBAS provides precision position services limited to use in the approach phase only. SBAS provides a precision position service in all flight phases, but GBAS provides better accuracy in the approach phase. On the other hand, the ABAS/ABIA system approach is particularly well suited to distinctively increase the levels of integrity and accuracy (as well as continuity in multi-sensor data fusion architectures) of GNSS in a variety of mission- and safety- critical applications. Synergies between these three GNSS augmentation systems (Fig. 36) can be studied to develop a Space-Ground-Aircraft Augmentation Network (SGAAN) that can be used for a number of safety-critical aviation applications (Fig. 37). Fig. 37. Space-ground-aircraft augmentation network.
Fig. 38. SGAAN integrity Augmentation architecture.
PMO techniques would have the benefit of providing a mathematical optimum. However, they can be used instead of Fig. 40. ABIA system architecture for UAS applications [148].
DCO only if the SAA time horizon allows (i.e., only if the time to conflict is much greater than the computational time required to obtain an op timal trajectory). This could be the case, for example, in properly developed air traffic separation maintenance systems. Error analysis is performed to determine the overall uncertainty volume in the airspace surrounding the intruder tracks based on the cooperative/non-cooperative unified method described in [146]. This is accomplished by considering both the navigation and the tracking errors affecting the measurements and translating them to uni fied range and bearing uncertainty descriptors. As discussed in the previous section, the SGAAN research demonstrated the potential of this new technology to enhance GNSS integrity performance in a variety of mission- and safety- critical applications including experimental flight test/flight inspection, precision approach and automatic landing (also in synergy with GBAS and SBAS) [148]. Therefore, an ABIA system concept was developed for UAS applications (Fig. 40). Fig. 39. SAA system architecture. A dapted from [146].
ABIA/SAA integrated architecture [148]. The key advantage 1s that the sate avoidance is determined by evaluating the risk-of-collision and then a safe manoeuvring point is identified from where the host UAV can manoeuvre safely (i.e., any manoeuvre can be performed within the UAV operational flight envelope). The risk of collision is evaluated by setting a threshold on the probability density function of a near mid-air collision event over the separation volume. The risk-of-collision is zero at the safe manoeuvring point. If both the safe-separation thresholds are violated, a mid-air collision threat is detected and the SAA WIF is generated. To prevent any WIF, the flight path optimization process starts when the first CIF is generated. PMO and DCO techniques are used to generate a new optimised trajectory free of any integrity degradations. Depending on the relationship between the available time-to-collision and the computation time required to generate the optimal trajectory, PMO or DCO solutions are applied. The optimised trajectory data are sent to the AFCS (and to the ground pilot) for execution of the avoidance manoeuvres. In the trajectory optimisation process, the aircraft 3DOF/6DOF model defines the dynamics constraints, while the satellite elevations and the aircraft heading rates are used as path constraints. Results from simulation case studies confirmed that the Integrity-Augmented SAA (IAS) solution is capable of performing high-integrity conflict detection and resolution when GNSS is used as the primary source of navigation data [148-150].
Contents University of Nottingham, Nottingham Geospatial Institute, Nottingham, NG7 2TU, United Kingdom ‘RMIT University, School of Engineering — Aerospace Engineering and Aviation Discipline, Bundoora, VIC 3083, Australia Roberto Sabatini!*, Terry Moore? and Subramanian Ramasamy! Global navigation satellite systems performance analysis and augmentation strategies in aviation
Cholesky factorization is used to reliably compress the information in the covariance matrix (C). This factorization produces 4x4 an upper triangular matrix (R), which can be used to reconstruct the relative covariance matrix as follows: Table 17. SUDRE Indicators and values in Message Type 27.
Table 18. GBAS Service Level Performance and Service Levels (adapted from [36)).
7 Introduction in the environment (e.g., elevated terrain and tall buildings when flying at low altitude) or by the aircraft itself during manoeuvring (e.g., aircraft wings and empennage masking the on-board G NSS antenna), ionospheric scintillation, Doppler shift, multipath, jamming and spurious satellite transmissions. Anyone of these phenomena can result in partial to total loss of tracking possible tracking errors, depending on the severity of the e and the receiver characteristics. Tracking errors, especial and ffect y if undetected by the receiver software, can result in large position errors. Partial loss of tracking results in geometry degrada which in tum affects position accuracy. Consequently, G tion, SS alone does not always provide adequate performance in mission- essential and safety-critical aviation applications where levels of accuracy and integrity are required. high
Table. 1. Tropospheric delay for varying elevation angle.
Table. 3. Typical UEE error budget (95%). Adapted from [31].
Table 2. GPS single frequency C/A code UERE budget. Adapted from [31].
Table 4. SPS SIS URE accuracy standards. A dapted from [31].
Table 5. DOP expressions. divided by its volume (i.e., the RSS of the reciprocals of the 4 altitudes of the tetrahedron) [32]. Therefore, we can write: Table 6. GNSS figure of merit.
The TTFF is commonly broken down into three more specific scenarios, as defined in the GPS equipment guide: Table 7. Aviation GNSS Signal-in-Space Performance Requirements [37-41].
Table 8. Nominal receiver GPS signal power and received C/No [43]. As an example, Table 8 shows the expected C/N, for a receiver tracking a GPS satellite at zenith given a typical noise power density of -205 dBW-Hz [75]. The values of C/N, listed in Table 14 should be compared against the values required to acquire and track GPS signals. These thresholds are heavily dependent on receiver design and for most commercial GNSS receivers they are in the order of 28-33 dB-Hz for acquisition and 25-30 dB-Hz to maintain tracking lock [42, 43]. Current generation avionics GNSS receivers, which are designed to operate in high dynamics conditions, typically exhibit better C/N> thresholds [44, 45].
Table 10. Possible DGNSS data link frequencies. An uplink can be a separate transmitter/receiver system or the DGNSS signals can be superimposed on a GPS-link L-band ranging signal. The uplink acts as a pseudo-satellite or “pseudolite” and delivers the ranging signal and DGNSS data via the RF section of the user receiver, much in the same way the GPS navigation message is transmitted. The advantages are that the additional ranging signal(s) can increase the availability of the position solution and decrease carrier-phase initialisation time. However, the RS and UR’s become more complex, and the system has a very short range (a few kilometres at the most). This is not only because of the line of sight restriction, but also the power must be kept low in order to avoid interference with the real satellite signals (i.e., the pseudolite can become a GPS jammer if it overpowers the GPS satellite signals). A downlink option is also possible from the users to the RS or other central collection point. In this case the differential solutions are al calculated at a central location. This is often the case for test range applications where precise vehicle tracking is desired but the information is not used aboard the vehicle. The downlink data can be position data plus the satellite tracked, or pseudorange and deltarange measurements, or it can be the raw GPS signals translated to an intermediate frequency. The translator method can often be the least expensive with respect to user equipment, and therefore is often used in munitions testing where the user equipment may be expendable.
Table 11. Classification of DGNSS techniques.
Since the RR noise and multipath errors are included in the differential corrections and become part of the user’s error budget (root-sum-squared with the user receiver noise and multipath errors), the receiver noise and multipath error components in the non-differential receiver can be lower than the correspondent error components experienced in the DGNSS implementation. Another type of error introduced in real-time DGNSS positioning systems is the data link’s “‘age of the corrections”. This error is introduced due to the latency of the transmitted corrections (i.e., the transmitted corrections of epoch ty arrive at the moving receiver at epoch. These corrections are not the correct ones, Table 13. DGNSS L1 pseudorange errors with increasing distance from RS.
Table 12. Error sources in DGNSS. The ephemeris error is effectively compensated unless it has quite a large out-of-range component (e.g., 1000 metres or more due to an error in a satellite navigation message). Even then, the error will be small if the distance between the reference receiver and user receiver is small. Finally, the satellite clock error is compensated as long as both reference and user receivers employ the same satellite clock correction data. As already mentioned, the correlation of the errors experienced at the RS and the user location is largely dependent on the distance between them. As the separation of the user from the RS increases, so does the probability of significant differing ionospheric and tropospheric conditions at the two sites. Similarly, the increasing separation also means that a different geometrical component of the ephemeris error is seen by the RR and UR. This is commonly referred to as “Spatial Decorrelation” of the ephemeris and atmospheric errors. In general, the errors are likely to remain highly correlated for users within 350 km of the RS. However, if the distance is greater than 250 km the user will obtain better results using correction models for ionospheric and tropospheric delay [45, 79]. Additionally, practical DGNSS systems are typically limited by the data link to an effective range of around 170 km. Table 13 shows the error budget determined for a SPS DGPS system with increasing distances from the RR [45].
Table 15. Evaluation of GIVEIi [41].
Table 19. GBAS Service Levels (adapted from [36]). The minimum service volume to support Category II precision approach or autoland with any minima is the same as the minimum Category II service volume, with the addition of the volume within +450 feet from the runway centreline extending from the threshold to the runway end, from 8 feet above the surface up to 100 feet. There are different metrics for the GBAS performance; one of these is the SIS performance. This is defined in terms of accuracy, integrity, continuity and availability of the service [65]; this performance refers to the output of the “fault- free” airborne user equipment. The interaction between airborne and ground systems is defined by a separation of responsibilities:
Table 20. SBAS vertical and horizontal alert limits [41]. 39]. LNAV/VNAV approaches us provided by either barometric altime ter or GBAS. Aircraft to generate integrity flags based on the SBAS VPL. Similarly, the SBAS HAL is the threshold used to generate integrity flags based on the SBAS HPL. For oceanic, en route, terminal, or avigation/V ertical Navigation (LNAV/V NAV) approach VAL is not defined by the WAAS MOPS and ICAO SAR e lateral guidance (556 m ateral limit) from GNSS and/or GBAS and vertical guidance Lateral that do not use GBAS for the vertical guidance portion mus VNAV-capable altimeters, which are typically integrated with moder flight directors and/or F (FMS). t have ight Management Systems
Table 21. Vertical Alert Limit [40]. Similarly, the Lateral Alert Limit (LAL) defines the threshold values of LPL and PLPL. The LAL values are listed in Table 22. The Final Approach Segment Vertical and Lateral Alert Limits (FASVAL and FASLAL) are listed in Table 23.
Table 23. FASVAL and FASLAL. Table 22. Lateral Alert Limit [40].
Based on these alert limits and limitations, the criteria for producing SBAS/GBAS CIFs and WIFs can be set and are listed below:
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HIGH ANGLE OF ATTACK AERODYNAMICS
by Sravankumar Kota
Fig. 8.10. Types of boundary conditions on the six outer surfaces of the compu- a space IJK resulting from the O-O type grid mapping (Rizzi and Eriksson 1984
Fig. 8.11. Definitions of the streamline coordinates and flow angles
Fig. 1.1. Classification of flow structures over slender sharp-edge delta wings
Fig. 2.2. The rolled-up tip vortices behind a rectangular wing — cross-section rear view (Bippes 1972)
sponding flow fields. As the angle of attack is further increased, secondary vortices, as well as more complicated separated flows, are observed. Fig. 2.3. The rolling up of the vortices from the side edge of a rectangular wing An example of such an intriguing three-dimensional separated flow structure is observed at low speeds on the low AR rectangular wing at high angles of attack, as illustrated in Fig. 2.4. At the high angles of attack, the flow becomes more “complicated”, as local three-dimensional separations cause the separated shear layers to roll up into vortex “spirals” that detach from the planar surface of the wing. The visualization technique utilizes an oil film which is forced to follow the skin friction lines on the surface by the flow friction forces. Regions of attached flow and various types of separation pattern are observed on the wing surface, as shown in Fig. 2.4. The skin friction line pattern, which includes several singular points, is the basis for a topological interpretation of such flows. In this case, the particular interpretation for this flow was proposed by Peake and Tobak SJINON\ *. i2l. . .2.. 2. 2k’... 2 ' kb fn TAY UN Oi—== = Yd ln ge nde
Fig. 2.5. Topological singularities in the skin friction lines observed on the surface of the low AR rectangular wing at high angles of attack, according to Peake and Tobak (1980)
In some respects, this rolled-up flow is a much simpler flow structure than that observed on the low AR rectangular wing at reasonably high angles of attack. Some early flow visualization photographs, illustrating the rolling- up of the leading edge vortices over delta wings with sharp edges starting at low angle of attack, are shown in Fig. 2.6. The separation of a vortex sheet from the sharp leading edges, already at small angles of attack, and the rolling-up of the vortex sheet at moderate angles of attack, are very well illustrated in this figure. The process of the generation of the leading edges rolled-up vortices and the associated flow field, including the appear- ance of the secondary vortices as the angle of attack increases, is described graphically in Fig. 2.7.
Fig. 2.8. Visualization of streamlines and skin friction lines on a 75° delta wing (Monnerie and Werle 1968)
Fig. 2.9. Visualization of the rolled-up vortices over a delta wing — top view an ross-section view (Werle 1958)
Fig. 2.10. Oil flow visualization on delta wings at various Reynold numbers — effects of laminar, transitional and turbulent flows (Carcaillet et al. 1986)
Fig. 2.12. Three-dimensional flow separation over a delta wing at high angles of attack (Werle 1974, Legendre 1965)
Fig. 2.13. Flow structure on a thick delta wing with rounded leading edges at moderate and high angles of attack (Werle 1958
which includes the separation zones and the rolled-up vortices, is symmetric and very similar to the symmetric flow over the delta wing in subsonic flow. As the angle of attack is increased to a = 17°, the vortex flow is “suddenly” deformed into an asymmetric flow structure. This asymmetric flow struc- ture is stable; i.e. at the specific angle of attack, the flow remains steady at the specific asymmetrical positions of the rolled up vortices. As the an- gle of attack is further increased, the flow asymmetry may increase or, at certain angles of attack, “suddenly” the flow structure may “jump” into an opposite asymmetry (in this case, the first jump from left to right occurs between a = 18° and a = 19°) as is well demonstrated in Fig. 2.14. The most intriguing aspect of this phenomenon is the fact that initial direction
Fig. 2.16. A cross-section rear view of the rolled-up vortices over the slender Ogive-Cylinder at high angles of attack (Werle 1974)
Fig. 2.15. The vortex structure on a slender Ogive-Cylinder body at high angles of attack (Fiechter 1969)
Fig. 6.10b. Variation of vortex position with increasing angles of attack (Brown and Michael 1955)
Fig. 2.19. Views at various cross-sections along the elongated body of the asym- metric vortices and the corresponding topological interpretations (Jorgensen 1977, Peake and Tobak 1980)
Fig. 2.20. Flow structure over a square cross-section body at various roll attitudes and at moderate angles of attack
Supersonic flows over elongated bodies at angles of attack also (like in the case of slender wings) generate rolled-up vortices in the flow behind the body, as seen in the schlieren photograph presented in Fig. 2.21. It is noted that at supersonic speed of M = 4 and at a 5° angle of attack, the boundary layer is already separated and rolled-up vortices are already observed. Fig. 2.21. Schlieren photo of the flow over a 10° blunted cone-cylinder with a flare at a = 5° and M = 4 (Peake and Tobak 1980)
b. Primary and secondary vortices separation and reattachment at a = 18° illustrated in Fig. 2.22
Fig. 2.25. Oil flow pattern over a cone-cylinder-flare body at small (Ericsson, Reding and Guenther 1969)
Fig. 2.26a. A three-dimensional schematic of flow structure over a cone-cylinder- flare body with foci type separation (Peake and Tobak 1980)
Fig. 2.26c. Conjectured skin friction lines superimposed over calculations (Peake and Tobak 1980) Fig. 2.26b. Calculated streamline di- rections on the surface of a cone- cylinder-flare configuration (Hung 1979)
a Oil flow pattern on the shuttle-orbiter at a = 0 (Ericsson and Reding 1976
Fig. 2.32. Loss of ailerons effectiveness at increasing angles of attack for a delta- canard fighter configuration (Skow 1983)
Fig. 2.33. Schematic of the symmetric and asymmetric separation of vortices on the slender nose of fighter aircraft
Fig. 2.36. The loss of lateral control effectiveness due to rudder deflections at increasing angles of attack (Skow 1983)
Fig. 2.35. Wind tunnel and flight test data illustrating the yawing moment re- versal at increasing angles of attack
Chapter 2. Description of Flows at High Angles of Attack
Fig. 2.40. Side view of the asymmetric vortex trajectories due to sideslip on a fighter aircraft showing the trajectories of the forebody and LEX vortices (Skow 1983)
Fig. 2.39. The asymmetric vortex structure on the concorde due to combined angles of attack and sideslip (Skow 1983)
Fig. 2.41. Top view of the trajectories of the asymmetric vortices due to sideslip of a fighter aircraft (Skow 1983)
Fig. 2.42. Loss of rudder effectiveness due to combined variations of angle of attack and sideslip (Skow 1983)
Fig. 2.43. Loss of aileron effectiveness at increasing angles of attack in the pres- ence of sideslip for a fighter aircraft (Skow 1983)
Fig. 2.44. Top view of the breakdown of the leading edges vortices on a delta wing — symmetric case, no sideslip (Werle 1960) It is well established that concentrated vortex filaments suddenly break down at certain flow conditions. A view of breakdown of vortices on the delta wing at increasing angles of attack is shown in Fig. 2.44. The break- down can occur symmetrically, as in Fig. 2.45a or asymmetrically, as a result of sideslip, as in Fig. 2.45b. A side view of the asymmetric vortex breakdown on an aircraft configuration is shown in Fig. 2.46.
Fig. 2.46. Side view of the trajectories of the asymmetric vortex breakdown on a fighter configuration (Erickson 1979)
Fig. 2.47. The rolling moment reversal due to asymmetric vortex breakdown at increasing angle of attack (Skow 1983)
b. Cross-section visualization and illustration drawing
b. Cross-section view visualization and illustration drawing
Fig. 2.50. Vortex structure on an oscillating delta wing during the upstroke and downstroke pitching motions (Gad-E]-Hak and Ho, C.M. 1985
Fig. 2.51. The hysteresis loop of the vertical position of the vortex over an os- cillating delta wing (Gad-El-Hak and Ho, C.M. 1985)
Fig. 2.54. Experimental skin friction line pattern in a turbulent boundary layer oblique shock interactions (Peake et al. 1970)
Fig. 2.55. Surface heat transfer rate and pressure profiles on an inclined fin (Scuderi 1978)
Fig. 2.56. Side and front views of the damage to the ventral fin of the X-15A-2 aircraft:due to shock impingement (courtesy of NASA)
Fig. 2.57. Heat transfer rates and oil flow on the surface of a 10° blunt cone at M=6 and a = 20° (Zakkay et al. 1972)
Fig. 2.58. Interpretation of the singularities in the skin friction line pattern near the separation point of the blunt cone (Peake and Tobak 1980)
3.2 Topological Concepts for the Analysis of Vortical Flows Fig. 3.1b. Attached flow pattern on an elongated body
Fig. 3.2a. Illustration of skin friction pattern on bodies with two nodes of at- tachment, one node of separation and one saddle point of attachment
Fig. 3.2b. Illustration of skin friction pattern on bodies with two nodal points of separation, one nodal point of attachment and one saddle point of separation
Fig. 3.3. The skin friction line and vortex line patterns for saddle points of at- tachment and separation Based on these observations, Peake and Tobak (1980) define the dis- tinction between attached and separated flows as follows: The skin friction pattern for attached flow cannot contain any saddle points; the skin fric- tion pattern for separated flow must contain at least one saddle point. This is a necessary condition for defining separation, however, the full defini- tion of the separation phenomena will be included in the discussion on the characteristics of the singular points in Section 3.2.4.
Fig. 3.4. The pattern of skin friction lines and vortex lines near singularities (Lighthill 1963)
Fig. 3.5. Separation line of the saddle point between two nodes of attachment (Lighthill 1963)
Fig. 4.1. Mathematical formulation of the flow (Hess 1975)
Fig. 4.2a. The path of integration for two sections on the wing Chapter 4. Linear Aerodynamics of Wings and Bodies
Fig. 4.2b. The path of integration for two sections on the wing with vorticity generation line at the trailing edge
Fig. 4.4a. Spanwise loading distribution calculated by the modified. Multhopp method as compared with experimental data (Lamar 1968)
Fig. 4.4b. Centers of pressure calculated by the modified Multhopp method as compared with experimental data (Lamar 1968)
Fig. 4.4c. Chordwise loading distributions calculated by the modified Multhopp method as compared with experimental data (Lamar 1968)
Fig. 4.5a. Schematic of division of two-dimensional and axisymmetric configura- tions (Hess 1975)
Fig. 4.6. Comparison of the surface source density method calculations with the experimentally measured pressure distributions on a radome mounted on a fuselage (Hess 1975)
Fig. 4.7. Comparison of experimental surface pressure distributions and lift co fficients with the calculations of the modified surface-source method (Hess 1975
Fig. 4.8. Distribution of vortices on a typical lifting configuration
a. Spanwise distribution of section lift coefficient
b. Chordwise pressure distribution at 25% span station
Fig. 4.10. Layout of wing planform for VLM calculations — axis system, panel elements and horseshoe vortices (Margason and Lamar 1971)
Fig. 4.11. The definitions of the geometry of a panel (Margason and Lamar 1971)
Fig. 4.12. Schematic of the division of a wing-body-tail configuration
The planar boundary conditions are applied to the lifting surfaces only and have identical formulations in versions B and C. Wing thickness effects are represented by linearly varying source distributions, and wing lifting effects-are represented by linearly varying vortex distributions. Both sin- gularities are assumed to lie in the plane of the wing. The source-strength distribution is set equal to the known chordwise slope of the thickness dis- tribution, while the “no flow” boundary condition at the control points determine the strengths of the vortex distribution. The equations for the boundary conditions in the planar case are
Fig. 4.14. Results of calculations of Cp with USSAERO B and C programs and comparison with experimental data for swept uncambered wing at M = 0.80 and a = 4° (Woodward 1980)
Fig. 4.15. Results of the calculation of Cp using USSAERO B and C programs, and comparison with experimental data for uncambered arrow wing at M = 2.01 and a = 4° (Woodward 1980)
Fig. 4.17. The streamlines for a doublet with axis normal to the velocity direction
Fig. 4.19. The skewed horseshoe vortex in supersonic flow
Fig. 4.20. Various panel divisions and position of vortices for the supersonic VLM (Miranda et al. 1987)
Fig. 4.21. The plane vortex wake with yaw and pitch for the supersonic linear VLM calculations (Miranda et al. 1987)
Fig. 4.22. The aerodynamic coefficients of arrow wings calculated by the super- sonic program and by the conical flow theory (Miranda et al. 1977)
Fig. 4.24. Networks of surface panels for an aircraft configuration (Carmichael and Erickson)
Fig. 4.25. Subinclined and superinclined panels (Carmichael and Erickson 1981)
Fig. 4.27. Schematic of the PAN AIR architecture (Carmichael and Erickson 1981)
Fig. 4.26. Geometry for calculation of the perturbation velocity
Fig. 4.28. Calculated drag coefficients for a flat-plate delta wing of AR=2 at M. = .53; S; is wing area, S2 and S3 are control volume surface areas
a. Panel representation of wing-body configuration
b. Lift coefficient variation as a function of a
Fig. 4.32. Chordwise pressure distributions on a swept back wing profile of the EET configuration at 7 = 0.55 and a = 4.31° (Margason et al. 1985
Fig. 4.36. Comparison of calculated lift curve slope by various panel methods with experimental data for 37° swept back wing (Margason et al. 1985)
Fig. 5.2. Position of vortex center of gravity for elliptic loading (Sprieter and Sacks 1951)
Fig. 5.1. Shape of vortex sheet for elliptic loading (Westwater 1935)
Fig. 5.3. The roll-up of the edge of a two-dimensional vortex sheet into a spiral (Kaden 1931)
Fig. 5.5. Geometric notations used for the cutoff distance evaluation The cutoff distance, r*, defined by Eq. 5.11, is dependent on two pa- rameters characterizing the vortex model, namely the strength of the in- fluencing vortex, I’, and the lengths of both segments (i.e. the integration step length). This makes the cutoff distance criterion sensitive to local con- ditions at each step of the iterative process, rather than an arbitrary choice of a single valued cutoff distance used over the whole flow field. The size
Chapter 5. Vortex Flows and the Rolled Up Vortex Wake To prevent the generation of exaggerated induced velocities due to the singularity of the potential flow vortex elements, a “cutoff distance” was introduced by Almosnino, Zorea and Rom (1978) and is used in various discrete vortex line methods. A relation for the determination of the “cutoff distance” was presented by Almosnino (1985). Here, the “cutoff distance” is used to determine the range from the vortex segment, within which a solid body rotation can be assumed for the influencing vortex. This “cutoff distance”, r*, is determined by the equation
Fig. 5.7. The rolled-up vortex wake calculated by the modified vortex lattice method with finite core vortices
Fig. 5.9. Triangular elements in the elliptic cross-section wake (Portnoy 1975)
Fig. 5.8. Thick wake cross-sections and system of coordinates (Portnoy 1975)
Fig. 5.10. The initial roll-up of a 6% thick elliptic wake (Portnoy 1975)
Fig.5.11. The shape of the 6% thick elliptic wake at ¢* = 0.0192 (Portnoy 1975)
Fig.6.1. The trailing vortex sheet models in the linear and nonlinear lifting surface theories
Fig. 6.2. Normal components of velocity and force in the cross plane (Betz 1935)
Fig. 6.3. Nonlinear lifting surface vortex model for low AR rectangular wing
Fig. 6.4. Comparison of experimental results with the lift coefficients calculated by the nonlinear lifting surface theory for low AR rectangular wings (Bollay 1937)
Fig. 6.5. The vortex model for the nonlinear lifting surface theory of Gersten (1961, 1963)
Fig. ye The functions C, and C2 for wings of various planform shapes (Gersten 1961
Fig. 6.9. The conformal transformation coordinates and notations
Fig. 6.11. Comparison of calculated and experimental results for the lift coeffi- cients on delta wings (Brown and Michael 1955)
Fig. 6.12. The coordinates of the delta wing with the spiral vortex sheet (Mangler and Smith 1959) Chapter 6. Nonlinear Aerodynamics of Wings and Bodies
Fig. 6.13. Comparison of the calculated normal force coefficients by the Mangler- Smith method with experimental data and with results of the Brown-Michael method (Mangler and Smith 1959)
Fig. 6.14. Variation of Kp as a function of AR for delta wings (Polhamus 1966)
Fig. 6.15. Polhamus leading-edge vortex lift flow model (Polhamus 1966)
Fig. 6.16. Variation of Ky as a function of the aspect ratio for sharp edge delta wings (Polhamus 1966)
Chapter 6. Nonlinear Aerodynamics of Wings and Bodies Fig. 6.17. Comparison of Polhamus theory with experimental data for sharp edge delta wings of AR=0.5 to 2.0
Fig. 6.18a. Comparison of calculated lift coefficient with experimental data fo. sharp edge delta wings of AR=0.5
Fig. 6.18b. Comparison of calculated lift coefficient with experimental data for sharp edge delta wings of AR=1.0
Fig. 6.18c. Comparison of calculated lift coefficient with experimental data for sharp edge delta wings of AR=1.5
Fig. 6.18d. Comparison of calculated lift coefficients with experimental data for sharp edge delta wings of AR=2.0
Fig. 7.5. The induced drag coefficient for a rectangular wing of AR=1 (Zorea and Rom, 1978)
Fig. 7.6. The rolling-up of the side vortices near the tip-of.the rectangular wing of AR=1 at various angles of attack (Zorea and Rom, 1978) Fig. 7.7. The development of the vortex wake behind the trailing edge of the rectangular wing of AR=1 at a = 20° (Zorea and Rom, 1978)
Fig. 7.10. Variation of the normal force and pitching moment coefficients for : 45° swept back AR=1 rectangular wing
Fig. 7.11. Comparison of the calculated and experimental lift coefficients for a rectangular wing of AR=1/30 at angles of attack up to 40° (Zorea and Rom 1978)
Fig. 7.12. The flow over a delta wing obtained by the progressive deformation of a rectangular wing as proposed by Rehbach (1973) imbedding the bound line of the horseshoe vortex and deciding on the po sition of the control point in a triangular panel presents a difficult problem An interesting approach was presented by Rehbach (1973) where the flow over the delta wing is obtained by calculating the flow over a series of wing: starting with a rectangular geometry and progressively deforming the shapé¢ by narrowing the leading-edge to a pointed apex, until a triangular geo. metry is obtained, as shown in Fig. 7.12. The triangular geometry obtained by this process determines the positions of the imbedded horseshoe vor- tices and the control points in each deformed panel, as seen in Fig. 7.12 which in the last stage shows a panelling pattern which corresponds to the trapezoidal panelling method of Fig. 7.2d.
Fig. 7.13. Calculated vortex trajectories over a delta wing of AR=1 (Gordon and Rom 1985)
Fig. 7.14. Flow visualization of the rolled-up vortices over a delta wing of AR=2 (Rehbach 1973)
Fig. 7.15a. Variation of the lift coefficient with Aspect Ratio for flat delta wings at a = 20° (Rom et al. 1976)
Fig. 7.15b. Variation of the pitching moment coefficient with Aspect Ratio for flat delta wings at a = 20° (Rom et al. 1976)
Fig. 7.16. The 12% thickness delta wing geometry (Gordon and Rom 1985) The effect of thickness of the delta wing can be included in the NLVLM calculation. Such calculations were performed for both flat zero thickness and 12% thickness delta wings of AR=1, and the results were compared with the experimental data of Peckham (1961), shown in Fig. 7.16. The calculated and experimental lift, drag and pitching moment coefficients are presented in Figs. 7.17a. 7.17b and 7.17c resnectivelv. The lift coefficient for the thick wing is found to be about 15% to 20% below that for the zero thickness delta wing, and the pitching mo- ment is about 10% more positive in the thick wing case. Examination of these results indicate that the NLVLM calculations accord well with the experimental data. Very good agreement is obtained for the lift and the induced drag coefficients for both the zero thickness and for the thick delta wings, while the experimental pitching moments are a few percent more
Fig. 7.17a. Lift coefficient variation with angle of attack for the 12% thick delta wing (Gordon and Rom 1985)
Fig. 7.17c. Pitching moment coefficient variation with angle of attack for the 12% thick delta wing (Gordon and Rom 1985)
Fig. 7.18. NLVLM calculation of the trajectories and strengths of the leading edge panel and the second row panel vortices indicating secondary vortex sepa- ration (Almosnino et al. 1978)
Fig. 7.20. Variation of the aerodynamic coefficients as a function of a for the wing-canard European model B at 6, = 0 (Rom et al. 1990)
Fig. 7.21. Pressure distribution on the wing of the European model B configu- ration at a = 15° (Rom et al. 1990)
Fig. 7.23. Comparison of the NLVLM calculated local velocity vectors and the experimental results at the cross plan X/C = 0.8 and @ = 8.7° (Top — Oelker and Hummel 1988, Bottom — Gordon 1990)
Fig. 7.24. Oil flow visualization on the surface of a 80°/60° strake-delta wing at low speed (Brennenstuhl and Hummel 1982)
Fig. 7.27. The pressure distributions on the upper surface of the strake-delta wing at X/C = 0.75 (Experiment — Brennenstuhl and Hummel 1982, NLVLM - Gordon 1990)
Fig. 7.30. Onset of symmetric vortices on Ogive-Cylinder bodies
Fig. 7.29. Vortex arrangements and control point positions in the panels on the surface of a sharp nose body (Rom and Almosnino 1980)
Fig. 7.31b. Axial location of separation for blunt-nosed bodies
Fig. 7.31a. Axial location of separation for sharp-nosed bodies
Fig. 7.32a. Vertical location of concentrated vortex over bodies
Fig. 7.33. Nondimensional strength of concentrated vortex over body
Fig. 7.32b. Horizontal location of concentrated vortex over bodies
Fig. 7.34. Transition from symmetric to asymmetric vortex separation over bo- dies
Fig. 7.35. Lines of separation on the Ogive-Cylinger body (Rom and Almosnino 1980)
Fig. 7.36. Calculated trajectories of vortices shed from the Ogive-Cylinder body at a = 15° — Side view (Rom and Almosnino 1980)
Fig. 7.37. The roll-up of the vortex trajectories over the Ogive-Cylinder body (Rom and Almosnino 1980)
Fig. 7.38. Circumferential pressure distributions over Ogive-Cylinder body at various angles of attack (Almosnino and Rom 1983)
Fig. 7.39. Pressure distribution on the upper centerline of the shuttle nose and canopy (Almosnino and Rom 1988)
Fig. 7.41a. Variation of the lift coefficient as a function of angle of attack for the shuttle configuration (Almosnino and Rom 1988)
Fig. 7.41b. Variation of the drag coefficient as a function of angle of attack of the shuttle configuration (Almosnino and Rom 1988)
Fig. 7.41c. Variation of the pitching moment as a function of the angle of attack for the shuttle configuration (Almosnino and Rom 1988)
Fig. 7.42. Asymmetric separation on an Ogive-Cylinder — Cross flow view (Al- mosnino 1985)
Fig. 7.44. Definition of the vortex sheet method flow field model
Fig. 7.45. Location of discrete singularity parameters and control points for var- ious network types (Johnson et al. 1980)
Fig. a0) Panelling of a Thin Delta Wing and Body Configuration (Johnson et al. 1980
Fig. 7.47. View of the networks for the calculations of the thin delta wing and body configuration (Johnson et al. 1980)
Fig. 7.48. The parameters defining the kinematics of the free and fed sheets (Johnson et al. 1980)
Fig. 7.50. Comparison of the measured and calculated lift coefficient variation as a function of Aspect Ratio for delta wings
Fig. 7.49. Panelling of delta wing with wake (Johnson et al. 1980)
Fig. 7.51a. Variation of lift coefficient with angle of attack
b. Variation of pitching moment as a function of lift coefficient
c. Variation of the induced drag coefficient as a function of lift coefficient Fig. 7.51. Comparison of the measured and calculated lift, induced drag and pitching moment coefficients variation with angle of attack for delta wing of AR=0.25
Fig. 7.54. Pressure distributions over a flat and cambered delta wing of AR=1.15 at a = 25° (Johnson et al. 1980)
Fig. 7.55. Comparison of the Measured Drag Polar on Flat and cambered AR=1.15 delta wings with the FVS method and NLVLM calculations
Fig. 7.57. Pressure difference distributions on a flat arrow wing (Johnson et al. 1980, Manro et al. 1976)
Fig. 8.7. Grid lines in symmetry plane, on wing/fuselage/wake surface and in longitudinal plane touching the wing tip (Volpe et. al 1987a)
a. Planar view of the upper half, grid dimensions: 65 x 11 x 25
There is no significance to any particular position for placing the bound vortex in the panel in supersonic flow; however, maintaining uniformity with the subsonic geometry the bound vortex is placed at the 1/4 chord and the control point is positioned at the center of the 3/4 chord line of the panel. From the tips of each bound vortex, two semi-infinite trailing vortices are added to complete the horseshoe vortex of that panel. The velocity components at every point in the flow field can be calculated by summing up the velocities induced by all the panels; for example the velocity component in the z direction, u(z,y,z) can be expressed by (similar expressions can be written for v, and w) In the linear supersonic VLM, like in the linear subsonic VLM, the vortex wake is assumed to separate from the trailing edge and remain a flat vortex sheet in the plane of the wing. Applying the method devised for the subsonic NLVLM to the supersonic nonlinear case; the vortices are allowed to separate from the leading, side and trailing edges and their trajectories are calculated by an iterative procedure in which both the trajectories of the free vortices and the strengths of the bound vortices are simultaneously calculated.
Fig. 7.58. Flow Chart for the supersonic NLVLM (Tennenbaum 1990)
The equation for the state of an ideal gas in thermodynamic equilib- rium can be written as The conservation of mass, momentum and energy for a compressible invis- cid fluid in thermodynamic equilibrium, can be represented by the partial differential equations — the Euler equations — which can be written in the following form
In general, the grid system for a complex aircraft configuration is com- posed of many blocks matched at their interfaces, each one containing its own division of structured or unstructured meshes, with the possibility o having adaptive meshes in certain regions of the flow field. The large num- ber of options and combinations for the grid generation methods available to the interested user are presented in the many publications listed in the references of this chapter. Once the grid is established over both the confi- guration surfaces and the flow field, it is necessary to determine the various methods of solutions of the Euler equations by using the appropriate nu- merical scheme.
Fig. 8.4. Surface grid over an aircraft configuration (Bocci 1990)
b. Plane through the wing Fig. 8.5. Grid lines in symmetry plane, on wing/fuselage/wake surface and in cross-sectional plane (Volpe et. al 1987a)
a. Plane at leading-edge of aircraft
Fig. 8.6. Grid lines in symmetry plane, on wing/fuselage/wake surface and on the shell around the aircraft (Volpe et. al 1987a)
he local streamline direction along the n coordinate direction, the matriz B] is split into [B] = [B+] + [B~] where Similarly, in order to distinguish directions of propagation along the m coordinate direction, the matrix [C] is split into
Fig. pr. The O-O grid around the 55° cropped-delta wing (Eriksson and Rizzi 1984
Fig. 8.13. Velocity vector plot in the 67.5% chord plane of the 55° cropped-delta wing for M = 0.7 at a = 10° (Eriksson and Rizzi 1984)
Fig. 8.15. Comparisons of pressure distributions calculated by Euler TEAM code and FVS method with experimental data (Raj et al. 1986)
Fig.8.14. The cross-plane velocity field for the cropped-delta wing calculated with the Euler TEAM code (Raj et al. 1986)
Fig. 8.16a. Variation of lift coefficients as a function of angles of attack for the cropped-delta wing with sharp leading-edges (Sirbaugh 1987)
Fig. 8.16b. Variation of pitching moment coefficients as a function of angles of attack for the Cropped-Delta wing with sharp leading-edges (Sirbaugh 1987)
Fig. 8.17a. Variation of lift coefficients as a function of angles of attack for the cropped-celta wing with rounded leading-edges (Sirbaugh 1987)
Fig. 8.18. Spanwise pressure distributions for the cropped-delta wing with rounded leading-edges at the 80% chord station and at a = 10° and M=0.4, 0.85, 1.2 (Sirbaugh 1987)
Fig. 8.19. Spanwise pressure distributions for the cropped-delta wing with rounded leading-edges at the 80% chord station and at a = 20° and M = 0.4 and 0.85 (Sirbaugh 1987)
Fig. 8.21. Steady-state total pressure contours for a 75° swept delta wing at M=1.2 (Lee and Batina 1991)
Fig. 8.20. Unstructured grid about a sharp-edge 75° swept delta wing (Lee and Batina 1991)
Fig. 8.22. Vortex trajectories (particle traces} over the Cropped-Delta wing at M = 0.85 and @ = 21° illustrating the vortex breakdown (Sirbaugh 1987)
Fig. 8.24. Rolled-up flow in the 80% chord cross plane of a delta wing calculated from the Euler equations at M = 0.85 and at a = 20° and a = 24° (Scherr and Das 1988)
Fig. 8.26. Comparison of the experimental and computed lift coefficients for the strake on and strake off configurations (Raj et al. 1986)
Fig. 8.25. Oblique view of the C-H grid about the Strake-Wing-Body configura- tion (Raj et al. 1986)
Fig. 8.29. Isomach contours on upper surface and in the cross-plane of a fighter aircraft (Volpe and Jameson 1988)
Fig. 8.30. Computed streamlines on Chine-Forebody configuration at flow M = 0.2 and a = 25° (Volpe and Jameson 1988)
Fig. 8.31. Hybrid Surface Grid for Wing-Body-Pylon-Nacelle (Bocci 1990)
Fig. 8.32. Surface grid and isomach lines on Hermes type space plane calculated by Finite Element method at M = 10 and a = 40° (Viviand et al. 1987)
Fig. 8.33. Surface grid on the missile configuration (Lijewski 1987)
Fig. 8.35. Comparison of the calculated and measured normal force and pitching moment coefficients on missile configuration at M = 0.95 (Lijewski 1987) ‘
Fig. 8.37. Top view of the velocity vector field above the fighter canard-wing configuration at M = 0.9 and a = 24° (Epstein et al. 1989)
Fig. 8.36. Delta wing-canard fighter configuration (Epstein et al. 1989)
Fig. 8.38. Velocity vector field at three cross-sections on the fighter aircraft wing- canard configuration at M = 0.9 and aw = 24° (Epstein et al. 1989)
Fig. 8.39. Variation of the lift and drag coefficients as a function of angle of attack for the fighter aircraft configuration at M = 0.9 (Epstein et al. 1989
The pressure is related to the conservative flow variables by the equa- tion of state, given by
Fig. 9.3. Behavior of F(y) on rays on a body at high angles of attack (Degani and Schiff 1986)
Fig. 9.4. Skin friction lines on the unwrapped 6:1 prolate spheroid at 10° angle of attack (Rosenfeld et. al 1988)
Fig. 9.6. Surface shear vector and oil flow pattern on the surface of Ellipsoid (Pan and Pulliam 1986)
ih Particle traces in the cross-plane above the ellipsoid (Pan and Pulliam 1986
Fig. 9.10. Surface flow pattern on a tangent-ogive cylinder at a = 20°, Rep = 0.8- 10° with 40 x 91 x 97 grid (Hartwich and Hall 1990)
Fig. 9.12. Helicity contours in the cross-section at z/D = 6, M = 0.2, a = 20° and Rep = 5.0- 10° (Degani et al. 1990)
Fig. 9.13. Experimental and computed circumferential surface pressure distribu- tions on tangent-ogive cylinder at a = 20°, Rep = 0.2- 10°, 40 x 73 x 65 grid points (Hartwich an Hall 1990)
Fig. 9.14. Comparison of Navier-Stokes calculations with experimental data of the normal and side forces variation with angle of attack on tangent-ogive at M =0.18, Rep = 0.16 - 10° (Deese et al. 1991) by an explicit, fourth order, Runge-Kutta time-marching, finite-volume nu. merical scheme. The calculations are performed for Rep = (0.16-0.6) - 10° M = 0.18 and at angle of attack varying from a = 30° to 70°. The re sults of the calculations at a = 30° indicated a symmetrical voriex flov over the tangent-ogive body. By increasing the angle of attack to 50° with- out introducing asymmetric perturbation, the computation results actually indicated an asymmetric vortex structure on the leeward side of the ogive instead of the expected symmetric structure. This unexpected asymmetrica solution is attributed to numerical asymmetry hidden in the grid generatio1 method and the cell volume calculation procedure used in this calculation It is shown that the perturbation required to produce asymmetry is abou five orders of magnitude smaller than the body diameter. This very smal perturbation explains the great sensitivity of the symmetrical vortical flow which is in agreement with the experimental results showing an asymmetri vortical flow for angles of attack in the range of 35° to 60°.
Chapter 9. Solutions of the Navier-Stokes Equations Fig. 9.15. Circumferential surface pressure coefficient distributions (Degani and Schiff 1983)
Fig.9.16. Pressure distributions on 7.5° cone at M = 2.94, a = 22.6° and Rez = 5.85- 10° (Degani and Schiff 1983)
Fig. 9.18. Pressure distribution on a 10° cone at M = 7.95 and Rez = 0.4: 10° (Lawrence et al. 1987)
Fig. 9.20. Crossflow streamline pattern computed on a 5° circular cone at a = 20° and Mach numbers 1.4 to 3 (Siclari 1990)
Fig. 9.19. Convergence behavior of an elliptic cone at M = 1.8 and a = 20° (Siclari 1990)
Fig. 9.21. Crossflow streamline patterns computed for three different cross sec- tions with 5° half angles and varying Aspect Ratios at M = 1.8 and a = 20° (Siclari 1990) 9.5 Solutions of the Navier-Stokes Equations
Fig. 9.22. Comparison of the calculated surface streamlines with experimental oil flow pattern on an AR=1 delta wing at a = 20.5°, Re=0.9 - 10° (Hartwich et al. 1988
Fig. 9.25. Comparison of the calculated and experimental variation of the lift ieee) with angle of attack on AR=1 delta wing, Re=0.9 - 10° (Hartwich et al. 1988
Fig. 9.26. Comparison of the vortex flow structure over the 80°/60° double delta wing at M = 0.3, Re = 1.3- 10° (Fujii and Schiff 1989)
Fig. 9.27. Variation of the lift coefficient with angle of attack for the 80°/60° double delta wing in incompressible flow, Re=1.3 - 10° (Hartwich 1988) sity contours. On the right side, the computed surface flow pattern and th leading-edge vortices are shown. A schematic of the crossflow pattern at tw representative locations, one on the strake and the other on the wing, is als sketched. The secondary separation is located approximately under the vot tex center on the strake section. On the wing section, both the se-paratio: and the attachment skin-friction lines which originate at the strake hav been moved further outward due to interaction with the vortical flowfiel caused by the wing vortex. At a = 10°, the strake and wing vortices remai distinct over the wing surface beyond the trailing-edge, while at 19°, th strake and wing vortices are seen to merge near the trailing-edge. Comparisons of the computed and experimental pressure distribution
Fig. 9.29. Global view of the flow pattern on the leeward side of the double delta wing at M = 0.22, Re=4- 10° (Ekaterinaris et al. 1991)
Fig. 9.30. Leeward side pressure distributions on the double delta wing at M = 0.22, Re=4- 10° (Ekaterinaris et al. 1991)
Fig. 9.31. Comparison of computed and experimental pressure coefficients on a 75° delta wing at M = 1.7, Re=3.56- 10° and at a = 4°, 12°, 20° (Thomas and Newsome 1986)
Vet See Pee ee Se ee Me MEM OM EE OULU GILLIS Fig. 9.32. Calculation of the hypersonic flow over a 75° delta wing at M = 10.17, Re=3.4- 10° and a = 15° (Ruffin and Murman 1988)
Fig. 9.33. Vortex trajectories over the AR=1 delta wing showing vortex break- er at a = 40°, Re=0.9- 10° with 65 x 101 x 81 grid points (Hartwich et al. 1988
Fig. 9.34. Vortex core and vortex breakdown region on a 75° swept delta wing at M = 0.3, Re=10° and a = 32° with 112 x 105 x 70 grid points (Ekaterinaris and Schiff 1990)
Fig. 9.35. Vortex core and vortex breakdown region on a 75° swept delta wing it M = 0.3, Re=10° and a = 40° with 112 x 105 x 70 grid points (Ekaterinaris and Schiff 1990)
Fig. 9.37a. Canard-wing-body flowfield grid topology (Tu 1991)
Fig. 9.39. Particle traces for the wing-body configuration with canard off and canard on at M = 0.9, Re=1.52- 10° and a = 12° (Tu 1991)
Fig. 9.38. Upper surface flow pattern for the wing-body configuration with and without canard for M = 0.9, Re=1.52- 10° and a = 12° (Tu 1991)
Fig. 9.40. Comparison of computed and experimental surface pressure coeffi- cients for the refined grid with and without canard at M = 0.9, Re=1.52 - 10° and a = 12° (Tu 1991) Chapter 9. Solutions of the Navier-Stokes Equations
Fig. 9.41. Comparison of computed and experimental forces and moment coef- ficients on the wing-body configuration with and without canard at M = 0.9, Re=1.52- 10° (Tu 1991)
Fig. 9.42. Top view of the F-18 HARV aircraft instrumented for surface pressure and oil flow (Rizk, Schiff and Gee 1990)
Computational results Flight test Fig. 9.44. Comparison of computed turbulent surface streamlines with flight test - side view, at M = 0.34, Re=13.5- 10° and a = 19° (Ghaffari et al. 1989)
Fig. 9.46. Surface grid of the F-18 (Rizk and Gee 1991)
Chapter 9. Solutions of the Navier-Stokes Equations
Fig. 9.49. Computed limiting surface streamlines on F-18 at M = 0.243, Re=11- 10° and a = 30.3° (Rizk and Gee 1991)
Fig. 9.51. Particle trajectories for flow over an F-15 wing-body-inlet configura- tion at M = 0.15, a = 20° (Kwon and Sankar 1991) Chapter 9. Solutions of the Navier-Stokes Equations
Fig. 9.52. Calculated axial pressure distribution for the FDL-5A configuration with various mass injection at M = 10.2, Re=2.1- 10°/ft and a = 5° (Narain, Muramoto and Lawrence 1991)
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Aircraft Performance Aircraft Performance. Maido Saarlas © 2007 John Wiley & Sons, Inc. ISBN: 978-0-470-04416-2 Aircraft Performance Maido Saarlas
by Puguh Schubby
3.1 FUNDAMENTAL PERFORMANCE EQUATION
The relationship between the lift coefficient and the flight velocity at 40,000 ft and at the maximum landing weight is given by
flight at 423 ft/sec at the point of stalling out the aircraft. In practice, for safety reasons, such low (stall) speed should never be attempted. Moreover, near-stall high lift coefficients may cause sufficiently high drag that can exceed the powerplant capability.
The engine data is somewhat simplified (see Appendix D), and is taken to be independent of flight velocity. The drag has been cal- culated from Eq. 3.24 with
By differentiating with respect to nondimensional velocity V, one cat recover the results already obtained in Subsection 3.3.2. The minimun drag occurs at
The answers follow in the same sequence.
Assuming that the pilot can maintain an indicated airspeed of 242 ft/sec, the plane can glide a distance of 10,000 _ _ . x= 0.0823 > 121,500 ft = 23 miles
As the last item in this example, the time to climb from sea level to 15,000 ft will be determined. Assume that the weight remains constant at 9,117 lb. The time to climb can be calculated, with best accuracy, from Eq. 4.5, by first plotting 1/v. = 1/(R/C) from Figure 4.5 or Table 4.2 the original data, as shown in Figure 4.3. Numerical evaluation of the area under the 1/v, curve gives
Eqs. 4.22 and 4.26 reduce for the case of B = O (thrust is independent of the velocity) to Eqs. 4.17 and 4.21, respectively.
fuel weight if the range is to remain constant. In Figure 5.6 an empty weight increase AW, (due to equipment, stores, etc.) indicates that an area AR has been lost. For same desired range an equal AR must be made available by an increase of W, by AW,. The primary reason for this disproportionate increase in fuel weight is the much higher specific range at the end of the flight (near W,) where lower aircraft weight requires less thrust power. For convenience, the results of integration of the Figure 5.6 can be displayed as a plot of range against the in- stantaneous weight, Figure 5.7, which permits a quick evaluation of the available range at any given weight and is a valuable tool in mission analweie
In Section 5.2 it was assumed that the specific energy (h + V?/2g remains constant during the cruise portion of the flight. Consequently the second term in the range equation vanishes. However, in derivins the range equation (e.g., Eq. 5.4), it was assumed that L/D is constant which, in turn, implies that aircraft attitude is constant. Subsequen sections have shown that either the velocity or altitude must ther change, with a resulting change in energy level. Usually, velocity 1: kept constant, and altitude variation brings about the resulting changes In order to estimate the magnitude of the omitted term, Rutowsk (see bibliography) replaced specific energy variation de with altitude dh and developed from the hydrostatic equation is obtained from logarithmic differentiation of the lift equation
Maximum endurance is obtained if the multiplier
Carrying out the differentiation and clearing the fractions, one finds
The quasi-steady analysis presented in previous chapters determines the general bounds of the aircraft performance: stall, maximum veloc- ity, ceiling, climb, range, and endurance. Although corrections were included for speed variations, the average steady-state nature of the phenomena permitted solutions without recourse to path-dependent in- tegration and without the need for direct consideration of accelerations. Take-off and landing are accelerated operations involving velocity var- iations from zero to safe flight speed, thus requiring a more exacting analysis of equations of motions. Although take-off and landing rep- resent only a small portion of the total operation of an aircraft, per- formance of these two phases is considered very important due to two
The effect of head wind on transition distances can be estimated by determining
It should be pointed out that the tan~' terms should not be added, as is often a customary procedure. There is a hidden singularity at V=1.
In a steady, unaccelerated flight all the forces acting on an aircraft are in equilibrium. For about 90 percent of flight, the simple equilibrium condition C,V* = constant holds, which governs simple and minor adjustments to the flight path. In point performance approach, this may be assumed even for modest rectilinear acceleration. This chapter con- siders flight where the path direction changes sufficiently rapidly so that the above condition may not be valid.
Figure 7.5 Turn in a Horizontal Plane
To estimate the time to turn through an angle y, it will be assumed that the flight path is a circle which then yields for time to turn through w radians
Thus, in absence of any other restrictions, one obtains
Eqs. 7.45 to 7.50 provide necessary tools for evaluating the minimum turn radius and, at the first glance, show that high thrust-weight ratio
which indicated a bank angle of 59°. The time to turn is obtained from Eq. 7.26
where the drag has been introduced from Eq. 3.21. Next, a parametric curve calculation is established for an arbitrary range of —500 = P, = 250 and a practical range of velocities with the load factor n as another parameter to assist in superimposing the P, curves in Fig. 7.6. The curve for P, = 0 signifies T = D, which provides the locus for sustained (steady) performance evaluation. Following the line for load factor n, = 5 to the P, = 0 curve (point B) gives the following:
Figure 7.10 A-4 Skyhawk Energy Maneuverability Diagram, Sea Level, W = 18,000 Ib
Figure 7.12 Climbing, Turning Flight Eqs. 7.68 through 7.70 are equivalent to Eqs. 7.20 through 7.24 of section 7.4 and are valid for climbing or descending flight (y < 0). For the case of flight without power (thrust), this set of equations can be developed further to yield some practical and interesting results. In earlier developments and examples it was shown that more power is needed in turns than in level flight. It will be shown now that as an airplane is descending in glide, the angle of glide will be increased if the plane is turning. In glide, as the plane turns through a heading of d, the height lost is
One of the more useful and productive tools in mission analysis and preliminary design is the so-called constraint plot. Usually it is found in the format where take-off thrust loading T,/W, is plotted against the take-off wing loading W,/S for mission requirements of interest. For a given set of mission specifications, called constraints, calculation of the range of typical T,/W, versus W,/S values defines the available solution space for that particular set of missions (see Figure 8.1). A plot of T, /W, against W,/S with all the applicable mission specifications as par- 924
endurance, and so on. Example 8.1 serves to illustrate intricacies and possibilities of this technique to focus on what is possible, what is needed, or what may be excessive.
Figure D.2_ AVCO Lycoming 10-540 Reciprocating Engine
Figure D.9 Pratt and Whitney JT9D-3 Turbofan Engine
Figure D.12 Allison T-56-A Turboprop Engine, Thrust
Figure D.13 Allison T-56-A Turboprop Engine, Specific Fuel Consumption
TABLE 3.1 Lift and Velocity Comparisons
TABLE 3.3 Velocity Comparisons, ft/sec; W = 12,000 Ib
TABLE 3.4 Example 3.4 Engine Data
TABLE 3.5 Propeller Efficiency, 7
TABLE 3.6 Maximum Velocities
TABLE 4.1 Climbing Conversion Factors
TABLE 4.2 Rate of Climb Calculation Procedure Although Figure 4.4 provides a quick overview of the climb per- formance at one altitude, it is limited and impractical because it is difficult to read the values with sufficient precision from the figure. In any case, the information is already contained in Table 4.2 in adequate detail. ey ee 2s in ox ee eT mores) eee ae
TABLE 5.1 Range Integration Method
TABLE 5.2 Comparison of Maximum Range and Endurance
TABLE 5.3 Summary of Range and Endurance Equations
TABLE 6.1 Average Coefficient of Friction Values
TABLE 6.2 Example 6.4 Data
TABLE 7.1 Data for Example 7.1
TABLE 8.1 Typical Mission Constraints The equations used to calculate the thrust loading as a function of the wing loading (or the reverse) are the same ones as used to evaluate performance of the individual missions in the point performance sense. However, since some of the expressions and procedures (e.g., take-off run) may be rather lengthy and involved and in order to gain quick overview, usually simplified equations are used. This also includes many other simplifying assumptions concerning the variation of the main parameters with altitude and velocity. The emphasis in this meth- odology is on trends and relative behavior rather than on precision.
TABLE 8.2 Standard Flight Path, Time b. The time along the flight path, from one altitude to another, is obtained by plotting 1/P, curves, at constant altitudes against /.. The numerical values have been computed already in a. The calcu- lations can be carried out by use of Eq. 8.9 in the difference form or the integral form. Here, both have been used. The results are shown in Figure 8.5. 1 an | 1 714° . v1 a © r |
TABLE 8.3 Optimal Path Points, Time
TABLE 8.4 Optimal Path Points, Fuel
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