AIAA White Paper on Fluid Dynamics Challenges in Flight mechanics


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AIAA White Paper on Fluid Dynamics Challenges in Flight mechanics

  1. 1. 47th AIAA Aerospace Sciences Meeting Including The New Horizons Forum and Aerospace Exposition AIAA 2009-744 5 - 8 January 2009, Orlando, Florida Taxonomy of Flight Mechanics Issues for Aircraft, and underlying Fluid Dynamics Phenomena Stephen C. McParlin* QinetiQ Ltd., Farnborough, Hampshire, GU14 0LX, UK Robert W. Tramel† DigitalFusion Solutions Inc., Huntsville, AL 35085 This White Paper has been produced by the Vehicle Aerodynamics Subcommittee of the AIAA Applied Aerodynamics Technical Committee, with contributions from interested parties inside and outside AIAA. The objective of this paper is to define a process by which the underlying fluid dynamics, that drive problems in aircraft flying and handling qualities, may be addressed in a systematic sense. This will enhance the confidence levels for broader use of CFD methods to the aerodynamic design of air vehicles, assessments of their aerodynamic characteristics and determination of the limits of their useable flight envelopes. The White Paper recommends a series of workshops aimed at addressing the ability of CFD methods to predict correctly the results of ‘building block’ experiments. This White Paper lists several significant fluid dynamics phenomena that are commended topics for future (both experimental and computational) workshops under the auspices of relevant AIAA Technical Committees. The paper proposes steps within the AIAA to arrive at such workshops. * Principal Aerodynamicist, Aerospace Consulting, A7 Bldg., Cody Technology Park, Associate Fellow AIAA † Staff Engineer/Scientist, Building 1, Suite 210, 5030 Bradford Drive, Senior Member AIAA 1 American Institute of Aeronautics and Astronautics Copyright © 2009 by QinetiQ Ltd. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.
  2. 2. Contents Contents 2 Nomenclature 3 I. Introduction 3 II. Flying Qualities Requirements 4 A. Certification standards 4 B. Design aspects of flying qualities 5 III. Flight Mechanics Phenomena (“Flight Mechanics 101”) 6 A. Longitudinal characteristics 6 B. Lateral characteristics 7 C. General characteristics of non-linear stability modes 7 IV. Taxonomy of Dominant Flow Physics and Flight Regimes for Aircraft 8 A. Low speed and subsonic flight 9 B. Transonic flight 9 C. Supersonic flight 10 V. Configuration Dependencies 10 A. Unswept/swept wing designs 11 B. Slender wing designs 12 C. Hybrid and non-slender wing designs 12 VI. Basis for the Taxonomy of Flight Mechanics Phenomena 13 A. Flight mechanics mode: 13 B. Flight regime: 13 C. Manoeuvre type: 14 D. Configuration type: 14 VII. Fluid Dynamics Phenomena Underlying Flight Mechanics Issues 14 A. Boundary-layer transition 14 B. Flow separation (no shock waves) 14 C. Shock-wave/boundary-layer interactions 15 D. Shock-wave/vortex interactions 15 E. Vortex stability, bursting and interactions 15 F. Mixed-flow regions: spanwise segmentation of attached and separated flows 15 G. Flow control 15 VIII. Causality – Which Fluid Phenomena? 16 A. Structure of the taxonomic matrix 16 B. Sources of data 16 IX. Discussion and Suggested Topics for Future Workshops 18 A. Transonic buffet 18 B. Flow separation from rounded leading edges 19 C. Flow separation upstream of the trailing edge 20 X. Summary and Proposed Next Steps 20 A. Summary 20 B. Proposed next steps 20 Acknowledgements 21 References 21 2 American Institute of Aeronautics and Astronautics
  3. 3. Nomenclature APA = Applied Aerodynamics AFM = Atmospheric Flight Mechanics ATR = Attained Turn Rate AWS = Abrupt Wing Stall BVR = Beyond Visual Range D = Drag DPW = Drag Prediction Workshop EASA = European Aviation Safety Agency FAA = Federal Airworthiness Administration FCS = Flight Control System FD = Fluid Dynamics L = Lift M = Mach number SPO = Short Period Oscillation STR = Sustained Turn Rate TC = Technical Committee UAV = Uninhabited Air Vehicle WVR = Within Visual Range = Turn rate I. Introduction I N recent years, the AIAA Applied Aerodynamics (APA) Technical Committee (TC) has sponsored a succession of CFD Drag Prediction Workshops (DPW), as a means of advancing the state-of-the-art in drag analysis for civil transport aircraft. The DPW series has been successful as a means of focusing attention on a specific aspect of industrial CFD use, in identifying technical issues in CFD application which need to be addressed, as a relative benchmarking activity for practitioners, and, not least, as a forum in which future research needs could be brought to the attention of appropriate funding sources. Based on the positive experience from the DPW series, the military aviation community within the APA TC considered the feasibility of a similar moment prediction workshop series. This did not come to fruition, for a variety of reasons, not least the availability of a suitable experimental data set upon which to base a CFD prediction workshop. In addition, there were concerns, based on a number of prior activities in this field, including the DPW series, that any failure to predict some aspects of flight mechanics with CFD might leave participants none the wiser as to the underlying causes. Issues have arisen in the DPW series, where specific fluid dynamics problems, such as development of trailing edge separations or localised separation bubbles, have not been solved by progressive global refinement of the meshes used for CFD1. Thus it appears that for some of these problems, the devil is in the detail, and it is unclear whether their intractable nature arises from deficiencies in modelling (e.g. time-averaging of unsteady effects, inadequate turbulence or transition models and deficient numerical dissipation schemes) or their discretization in space and time. Experience from the DPW series, based on relatively ‘clean’ wind tunnel models, points towards the use of successively denser meshes to spatially resolve local fluid dynamics issues. The feasibility and affordability of this approach for more complex configurations, at flight Reynolds numbers, are therefore questionable without the continued operation of Moore’s law, which suggests that, with the passage of time, more computing power for the same price will cure all ills. There are other challenges in the world of military aircraft. The relative maturity of CFD for these applications is somewhat less than in the world of civil transport design. The range of operating conditions, and configurations, for combat aircraft in manoeuvring flight is much more diverse than for the civil aircraft cruise cases considered within the DPW series. It is also true that the range of configurations and conditions encountered in weapons flight is yet more diverse still. Given the issues identified within the DPW series, and the increased scope for local problems on complex military aircraft configurations, the global mesh refinement approach might prove to be a costly way to proceed, with no guarantee of success. It was unlikely that a ‘one-size-fits-all’ approach would cover all areas of the military aircraft flight envelope in which it is important to predict flight stability characteristics, hence an alternative approach was considered. This is based on a building-block approach for specific fluid dynamics phenomena, to 3 American Institute of Aeronautics and Astronautics
  4. 4. determine the most appropriate level of modelling for each, through a series of workshops combining analysis of experimental data and evaluation of CFD methods. The adoption of a building-block approach to maturing CFD also recognises that, in many cases, the underlying flow physics driving problems in military and civil aircraft are similar, and that workshops, or a series of workshops, should not necessarily be predicated on whether the end objective was application to a civil or military application. It is also the case that building-block approaches are not, in themselves, particularly constrained by releasability issues, and thus are open to a broader range of participants from industry and academia. This is not the first time that a ‘building-block’ approach has been advocated as a means for validation of component technologies in CFD. Indeed, previous consideration of these from a bottom-up perspective has led to similar approaches being advocated over a period of many years, notably through the NATO Research and Technology Organisation2. This approach is more considered and methodical than the alternative of considering a complex configuration problem in isolation, but it also offers a more thorough means of identifying specific issues that need addressing in CFD toolsets, hence providing the knowledge which might not arise from the alternative approach. It is the case that the flight mechanics problems encountered by aircraft are, with the exception of those specifically determined by their inertia properties, a consequence of the aerodynamic loads to which they are subject. Through the simple application of Newtonian mechanics, any change in the stability of an aircraft, whether progressive or abrupt, has a root cause in the change of distribution of aerodynamic load somewhere on its surface. Hence non-linear stability characteristics are usually attributable to changes in flow topology somewhere in the aircraft flow field. Structural elasticity under changing aerodynamic load may contribute to the non-linearity. To be able to predict the onset of non-linear stability characteristics with CFD therefore requires that the method should be able to predict the specific fluid phenomenon that provides causation. It is thus proposed that the means by which to address prediction of non-linear flight stability characteristics with CFD is to focus effort on prediction of specific fluid mechanics phenomena relevant to vehicle configurations under consideration. The objective is to identify through taxonomy, for particular flight mechanics problems, the causative fluid dynamics phenomena, then determine, through a programme of workshops, the most appropriate level of modelling to capture the flow physics adequately. These workshops should seek to separate out model deficiencies from numerical issues. Where successful, these workshops should disseminate best practice for predicting these phenomena in terms of both the level of physical modelling and discretization of the problem. Where unsuccessful in predicting a particular phenomenon, appropriate levels of interest should be stimulated. II. Flying Qualities Requirements A. Certification standards Flying qualities are a safety and qualification issue, as formally addressed in certification rules and standards. They are also a multidisciplinary issue, where aerodynamic design, flight control system design, structural design and, where necessary and appropriate, flow control, can all make a contribution to a successful outcome. We are concerned here with the proportions of the mix, particularly where it is not clear that the outcome will be successful. All of these aspects of the engineering solution have their limits, which should be understood in the context of the formal certification requirements. This is expressed in outline in the relevant UK Defence Standard3: “It is desirable that the specified flying qualities should be achieved by good aerodynamic and mechanical design. However automatic devices may be used where an overall benefit accrues provided that the system as a whole meets the requirements.” Further to this, the equivalent US Military Standard for piloted aircraft, MIL-STD-1797A4 states: “The aircraft shall be…resistant to departure from controlled flight, post-stall gyrations and spins. Adequate warning of approach to departure shall be provided. The aircraft shall exhibit no uncommanded motion which cannot be arrested promptly by simple application of pilot control.” These certification standards, as is appropriate, have a good deal of commonality in their treatment of flying qualities. Both contain more detailed qualitative and quantitative requirements and draw on a common basis for rating flying qualities: the Cooper-Harper ratings scale5, as shown in Figure 1. The Cooper-Harper rating process was intended for use on piloted aircraft, rather than their uninhabited counterparts. The subjectivity inherent in pilot- based assessment is handled by means of having a sufficiently large sample of pilot opinion. The principle embodied in a Cooper-Harper rating 1-2, that pilot compensation is not a factor in achieving desired performance, is directly applicable to uninhabited air vehicles (UAVs), where autonomous flight, without remote pilot intervention, is a specific design objective. The certification standards for military aircraft are written for a wide variety of classes of air vehicle, from light aircraft through to large multi-engine transport aircraft and surveillance platforms, as well as aircraft required to 4 American Institute of Aeronautics and Astronautics
  5. 5. perform manoeuvres over a large variety of operating conditions. However, the certification of light aircraft types and transports for non-military uses is governed by the appropriate civil authorities. Within the EU these are the responsibility of the European Aviation Safety Agency (EASA) and in the US, the Federal Aviation Administration. As with most international standards, these are harmonised between states and are thus broadly comparable. The European standard for Large Aeroplanes (CS-25)6 has a number of specific requirements relating to flying qualities, With regard to stall characteristics: “It must be possible to produce and to correct roll and yaw by unreversed use of aileron and rudder controls, up to the time the aeroplane is stalled. No abnormal nose-up pitching may occur. The longitudinal control force must be positive up to and throughout the stall. In addition, it must be possible to promptly prevent stalling and to recover from a stall by normal use of the controls” (CS 25.203). Likewise, for stall warning: “Stall warning with sufficient margin to prevent inadvertent stalling with the flaps and landing gear in any normal position must be clear and distinctive to the pilot in straight and turning flight” (CS 25.207). In the case of buffeting: “The aeroplane must be demonstrated in flight to be free from any vibration and buffeting that would prevent continued safe flight in any likely operating condition” (CS 25.251). Hence, the underlying principles for civil and military certification are: that stability characteristics should be consistent and predictable; that control departure boundaries should be well-defined, approached with adequate levels of warning and easily recoverable from. The primary risks of failure are therefore associated with abrupt and/or unpredictable changes in stability, and the lack of control power to overcome these. 1. Excellent – highly desirable. Pilot compensation not a factor for desired performance. 2. Good - negligible deficiencies: Pilot compensation not a factor for desired performance. 3. Fair – some mildly unpleasant deficiencies: Minimal Yes pilot compensation required for desired performance. 4. Minor but annoying deficiencies: Desired performance Is it No requires moderate pilot compensation. satisfactory Deficiencies warrant without improvement improvement? 5. Moderately objectionable deficiencies: Adequate performance requires considerable pilot compensation. Yes 6. Very objectionable but tolerable deficiencies: Adequate performance requires extensive pilot compensation. Is adequate performance No Deficiencies require attainable 7. Major deficiencies: Adequate performance not improvement with a tolerable attainable with maximum tolerable pilot compensation. pilot workload? Controllability not in question. 8. Major deficiencies: Considerable pilot compensation is required for control. Yes No 9. Major deficiencies: Intense pilot compensation is Is it Improvement required to retain control. controllable? mandatory 10. Major deficiencies: Control will be lost during some portion of required operation. Pilot decisions Figure 1 - Cooper-Harper Handling Qualities scale (from NASA TN-D 5153) B. Design aspects of flying qualities As expressed in the UK Defence Standard, there are some cost and robustness advantages in achieving desirable stability and control characteristics by the simplest passive means available, without recourse to complexity, either in mechanical systems or software. However, where increased performance demands and complexities in the underlying aircraft stability characteristics arise, modern flight control systems offer a means of overcoming deficiencies in the natural response of the vehicle. We are not primarily concerned here with the details of flight control system (FCS) design per se, but from the air vehicle system perspective, we must be aware of the issues which make FCS design more difficult and risky: 5 American Institute of Aeronautics and Astronautics
  6. 6. 1) Non-linearity of aircraft open-loop response with respect to variations in attitude. 2) Non-linearity of aircraft response with respect to control inceptor inputs. 3) Lack of control power available from control effectors. Of these, the last is the most fundamental. With sufficient control power and actuation bandwidth, it is perfectly feasible to render highly unstable open-loop responses controllable. Without sufficient control power, simple, stable, linear, systems will not respond adequately to inceptor demand. Availability of control power is thus an issue for initial configuration layout, rather than the details of aerodynamic design and flow development. The first and second of the listed items are of more interest in the context of aerodynamics. Non-linearity of aircraft response, when subjected to either a steady-state condition or a dynamic manoeuvre, is primarily a consequence of a change in aerodynamic loads, although inertia loads have historically been responsible for control departure of specific classes of air vehicle7. Non-linearity of control response is frequently associated with variations in the aerodynamic load associated with the associated effectors, particularly where these are close-coupled and subject to significant interference effects, such as near-field wakes. Aeroservoelastic effects are also often an underlying issue in non- linear control response, and these are generally treated by modification to the associated structure and mass properties, rather than through aerodynamic fixes. It is apparent that, at the initial stages of a project, the primary influence of aerodynamic design on flight control risks is the ability to eliminate non-linear responses to changes in aircraft attitude within the desired flight envelope, or delay them to more extreme flight conditions. Hence, if designing with flying qualities as an explicit or implicit design objective, the principles of consistency, predictability and avoidance of non-linear aerodynamic phenomena should be borne in mind. However, as with flight control system design, where system performance demands require compromises, it is not always possible to achieve desired flying qualities through aerodynamic design alone. An aerodynamic design based on consideration of the required flying qualities, however, remains the basis for all subsequent palliatives and improvements derived from flight control systems and any flow control fixes. The better the starting point from an aerodynamic standpoint, the less complexity will be required in the flight control system, while hopefully, the need for retrospective (and potentially compromising) fixes will be still further reduced. III. Flight Mechanics Phenomena (“Flight Mechanics 101”) From the standpoint of developing taxonomy of flight mechanics issues, it is generally appropriate to consider the classical linear stability modes, derived from small-disturbance analysis of the linearised equations of motion, before considering how these modes can develop non-linear aspects. The development of non-linear stability modes, including divergent and limit cycle behaviour, forms the basis for the majority of problems in flight control that we are seeking to address. Note that, given sufficient control power and actuation bandwidth, an unstable, yet linear, system is not particularly problematic. Difficulties are more likely to arise when the underlying assumptions of linearity break down. In some cases, the non-linear stability modes which arise may be seen as variations of the linear stability modes, particularly if periodic aerodynamic forcing occurs close to the frequencies associated with oscillatory modes. However, there are any number of purely non-linear modes which can occur as the flight envelope required of combat aircraft expands. Coupling between modes is also more likely at extreme flight conditions. The situation with regard to civil transport aircraft is much more clearly defined. Safety is the overriding issue, and hence avoidance of non-linear phenomena, including those arising from aeroelastics and aeroservoelastics, becomes a design objective. Determining the margin to onset of these phenomena from normal operating conditions then becomes a design issue. A. Longitudinal characteristics The numerous reference texts covering the area of classical flight mechanics describe the longitudinal and lateral stability response to a small disturbance in terms of quartic equations. The roots of the stick-fixed longitudinal quartic are two complex conjugate pairs, representing two distinct oscillatory modes in pitch and heave. These are conventionally described as the Short Period Oscillation (SPO) and Phugoid mode respectively. For a conventional, longitudinally stable aircraft, the SPO is a heavily damped oscillation of relatively short duration and high frequency, while the Phugoid mode is a lightly damped oscillation of much lower frequency. However, the performance benefits, in terms of reduced trim drag and increased agility, associated with reduced static margin make it likely that both SPO and Phugoid modes can be unstable in their open-loop form, requiring automatic stabilization through feedback control. Meeting the overall system requirements then requires more effort in the design and flight clearance of appropriate control laws. 6 American Institute of Aeronautics and Astronautics
  7. 7. Moving on from the linear stability modes, there are a number of commonly experienced non-linear modes in pitch. These correspond to forward or aft shifts in the location of the aerodynamic centre, resulting in pitch-up, tuck- under or unsteady variants of these. The divergent pitch-up and tuck-under modes can usually be seen in static loads obtained from wind-tunnel tests, while the oscillatory versions of these are generally the consequence of the associated variation of static margin feeding through into variants of the existing linear modes. Large-amplitude pitch oscillations frequently develop hysteresis loops, further complicating their representation. Hence, more accurate, non-linear representations of the longitudinal modes should incorporate higher-order modelling than the first order linear stability derivatives to capture this behaviour adequately. Uncontrollable pitch-up can and does have significant consequences for the useable flight envelope of an aircraft, particularly when at low speed and high lift. There are very many examples of aircraft subject to this particular problem, and, in practice, the usual solution is to limit the useable angle of incidence. The underlying causes of pitch-up are well known and understood, but are sensitive to flight regime and air vehicle configuration. Indeed, pitch-up at low speed and high speed can have similar symptoms, but different driving mechanisms. Tuck- under is a less usual problem, and is the mirror image of pitch-up, where divergence in pitch occurs at low and negative angles of incidence. Both pitch-up and tuck-under are associated with a forward movement of the aerodynamic centre, and hence decreased static margin. There is also the specific mode of Mach tuck, associated with the rearward movement of the aerodynamic centre with Mach number. This is of particular note at or near the transonic drag rise, resulting in a significant increase in control power requirements and trim drag to achieve a given manoeuvre. In effect, this becomes a subtle limit on the useable angle of incidence at transonic and supersonic conditions, as the control power required to trim increases beyond that available. B. Lateral characteristics By way of contrast with the longitudinal stability quartic, the classical lateral-directional stability quartic has two real roots and two complex conjugate roots. These produce two steady modes, the spiral and roll subsidence modes, and one oscillatory mode, the Dutch roll mode. The spiral mode is a tendency to simultaneously turn and bank. Although there are numerous cases of aircraft with an unstable spiral mode, the rate of divergence in these cases tends to be sufficiently slow to be easily controllable by manual means. The roll subsidence mode is, likewise, a slow convergence or divergence in roll. Like the spiral, this is a weak mode in most practical cases. The Dutch roll mode is more significant than the two others, being an oscillatory coupling between motions about the roll and yaw axes. Notably, the stability of the Dutch roll and spiral modes interacts. In general, increases in the rolling moment due to sideslip tend to promote instability in the Dutch roll mode, while having the opposite effect on the spiral mode. The converse is true of the yawing moment due to sideslip. As the Dutch roll mode is of relatively greater importance in flying qualities, this is generally given precedence over the spiral mode. As with the longitudinal modes, there are some common non-linear lateral modes, which in some cases can be related to the underlying linear modes, although at more extreme flight conditions, purely non-linear and coupled modes can develop. The divergent modes about the roll and yaw axes, wing drop and nose slice (or yaw off) respectively, are analogous to pitch-up/tuck-under. These represent static instability about the respective axes. Both of these phenomena can occur at nominally zero sideslip when driven by asymmetry in the flow development. As variations in static stability characteristics feed into the dynamic modes, the Dutch roll and spiral modes can be impacted by non-linearity. In particular, the oscillatory wing rock mode evolves from the Dutch roll mode when aerodynamic damping terms fall to zero or become negative, with peak response being shown for specific, narrow frequency bands. C. General characteristics of non-linear stability modes In general, at the boundaries between the linear and non-linear stability modes, the non-linear modes represent a variant of the linear modes in which one or more of the underlying stability derivatives exhibits non-linear characteristics. The risk to a successful engineering solution providing acceptable flying qualities is determined by the extent of the non-linearity and the rapidity of onset. Hence, at this point, it is desirable to address the five specific modes which are likely to have the most impact, ordered by flight mechanics axis: 1) Wing drop, defined as a divergence about the roll axis. 2) Pitch up/Tuck under, defined as a divergence about the pitch axis. 3) Nose slice, defined as a divergence about the yaw axis. 4) Pitch oscillation, driven by non-linear forcing of the SPO or Phugoid modes. 5) Wing rock, driven by non-linear forcing of the Dutch roll mode. 7 American Institute of Aeronautics and Astronautics
  8. 8. This list is by no means comprehensive. The classical linearised stability quartics are based on a number of simplifications and assumptions which do not hold across the full extent of the flight envelope, and hence the full non-linear equations of motion are more appropriate for modelling the flight dynamics of agile air vehicles. IV. Taxonomy of Dominant Flow Physics and Flight Regimes for Aircraft As part of their typical operating conditions, combat aircraft are required to operate over a much wider range of conditions, in terms of attitude, altitude and Mach number, than civil transport aircraft. Mission requirements can range from very high lift coefficients for launch and recovery at low Mach number through to very low lift coefficients at the opposite end of the Mach number envelope and low altitude. Manoeuvre conditions for modern combat aircraft are now much broader than the traditional Within-Visual-Range (WVR) transonic sustained manoeuvre case, and can include both Beyond-Visual-Range (BVR) combat at supersonic Mach numbers and post- stall manoeuvres beyond the normal lift limits at low speed. Each operating regime is subject to very different dominant flow physics, while the non-linear stability modes discussed above can be driven by completely different flow mechanisms at different flight conditions. It is thus worth considering the varying nature of the flow fields at different operating conditions, and the implications for developing taxonomy in which operating conditions are a factor. Figure 2, from ref. 8, shows a plot of turn rate, , vs. Mach number for three supersonic combat aircraft with representative combat mass at the tropopause. Note that the peak sustained turn rate (STR) is usually, but not always, at high subsonic Mach number, depending on the design requirements for a particular aircraft. The STR for each of the aircraft shown in fig. 2 is thrust-limited at this altitude and significantly below the structural limit, which might apply at lower altitude. Peak STR is thus determined by the combination of propulsion system and aircraft drag. Increasing thrust/weight ratio or reducing drag will improve the STR of the aircraft. Figure 2 – Sustained Turn Rate vs. Mach number for a set of combat aircraft at the tropopause (from ref. 8) The requirements for civil transport aircraft involve a more limited number of operating conditions. However, the requirements for safety, passenger comfort and fuel economy are more stringent at these conditions. The design drivers are thus different, although many issues associated with handling qualities are common. It is perfectly acceptable (and desirable) to perform manoeuvres with a combat aircraft which would not be the case with a civil transport aircraft. 8 American Institute of Aeronautics and Astronautics
  9. 9. A. Low speed and subsonic flight At low speeds, dynamic pressure, and therefore the control power available from aerodynamic control effectors, is at a minimum. Hence it is usual that the size of control effectors, and the response time of actuators, is driven by low-speed requirements. Operations below minimum drag speed also offer the prospect of speed instability, and hence throttle response becomes an issue. For both civil transport and combat aircraft, low speed operations revolve around launch and recovery, and for combat aircraft, to a more limited extent, low speed manoeuvre. High lift is a factor in all of these requirements, but to different extents for each. Fixed-wing aircraft, with some notable exceptions, are constrained to be launched and recovered horizontally, by means of a runway, occasionally with assisted takeoff and recovery. Takeoff requires acceleration to rotation speed at maximum mass within a specified distance and hence drag is an important factor. It is therefore desirable that the flow over the aircraft at these conditions is close to fully attached, to minimise drag during acceleration to rotation speed, and subsequently, on climb-out. Thus take-off is generally the least demanding low speed flight condition in terms of flow complexity, although it is frequently the case that control effectors in pitch are sized by the need to lift the nose wheel on rotation. To recover an aircraft within the constraints of a required field length, or on the deck of an aircraft carrier, drives minimum speed, and hence maximum lift coefficient. Issues of tail scrape and pilot visibility limit the useable angle of incidence at these conditions, and hence drive wing area and/or the complexity of the high lift system required. Civil and military airworthiness requirements include specific requirements for crosswind landings; hence at landing speeds there is a requirement to be able to sustain a trimmed angle of sideslip, as well as incidence. Drag is much less of an issue during landing than at takeoff and hence it is not unusual for the flow to be separated at landing. Flow unsteadiness, or buffet, at the approach condition, though, is highly undesirable. The flows at this flight condition therefore tend to be steady, separating or separated. In addition to the launch and recovery cases, for agile combat aircraft, there is the issue of low-speed manoeuvre. Unlike the landing conditions, these are not constrained by arbitrary limits on angle of incidence. These are characterised by two different flight regimes: lift-limited, where manoeuvre occurs up to maximum useable lift, and post-stall manoeuvre, where angular rates are still being generated beyond maximum lift, using alternative control effectors, such as vectored thrust. These manoeuvres tend to occur at Mach numbers below the corner point, at which attained turn rate (ATR) is a maximum. They are thus low-energy, transient manoeuvres, characterised by high angular rates, therefore the onset flow conditions are changing rapidly. The flows resulting at these conditions are highly separated and unsteady. Pitch-down recovery from these conditions competes with the nose wheel lift requirement to size the primary control effectors in pitch. B. Transonic flight Like large subsonic transport aircraft, combat aircraft spend the majority of their flight time in the high subsonic to transonic regime, with most time being spent cruising at optimum range (maximum ML/D) or loiter (maximum L/D). Operation at constant equivalent air speed means that increasing altitude drives towards higher Mach number. Combat is generally of short duration, but typically will require acceleration from a cruise or loiter condition to engage, followed by either WVR or BVR engagement, in which turns are made and conservation of energy is a major factor. In WVR combat, maximising turn rate, while minimising energy loss, drives towards aircraft operating at or near their peak STR for a given altitude. Potential energy can be traded for kinetic energy by losing altitude, with disengagement becoming an issue as altitude, and hence Mach number, reduces. Excursions above the STR limit allow tighter turn radius, and thus potentially target acquisition and engagement, at the cost of energy bleed-off and subsequent deceleration. Typical Mach numbers for maximum STR at the tropopause tend to be in the high subsonic regime. Openly published values8 of Mach number for peak STR at the tropopause for the F-16, MiG-21 and F-4 are approximately 0.7, 0.8 and 1.2, respectively, as shown in Figure 2. The transonic sustained turn regime has typically been a primary design point for combat aircraft with a primary air-superiority role. Drag minimisation at moderate to high lift coefficient is a design objective here, with the design lift coefficient at sustained manoeuvre being typically over 4 times that for cruise at the same Mach number and altitude. The vast majority of the drag at these conditions is therefore lift-dependent. Hence, avoidance of flow separation is desirable, but not always feasible. At high subsonic manoeuvre cases, compressibility has a significant impact on the nature of the flow and on flow separation mechanisms. Strong shock waves occur and, depending on the details of the configuration, become the primary sources of both lift-dependent drag and flow separation as Mach number increases. Flows at transonic manoeuvre conditions therefore tend to be separating to separated, with embedded shocks and consequent steady 9 American Institute of Aeronautics and Astronautics
  10. 10. and unsteady flow separations. Manoeuvre limits can also be imposed by structural strength considerations, particularly where there is a significant level of unsteadiness in the flow. For civil transport aircraft powered by turbofans, maximising range for a given fuel load, as dictated by the Breguet range equation, requires flight at conditions near or slightly into the transonic drag rise. However, the requirements for avoidance of buffet (measured at the pilot’s station in the cockpit) are more stringent than for combat aircraft. This means that the onset of shock-induced flow separation, or trailing edge separation downstream of a shock, is a condition to be avoided in normal flight envelope operations, including moderate manoeuvres, as defined in airworthiness requirements. A civil transport aircraft operating at these conditions will generally tend to have a wing loading somewhat higher than a combat aircraft operating at the same altitude and Mach number. The civil transport aircraft will thus normally have less buffet margin at transonic cruise than a combat aircraft. Determination of buffet margin thus becomes a major issue in the aerodynamic design of civil transports. C. Supersonic flight Until relatively recently, combat aircraft would only be expected to accelerate to supersonic conditions for relatively short-duration dashes, to either prosecute a time-sensitive target or evade an engagement. Upon commencing a manoeuvre, the aircraft would decelerate along the thrust-limited turn locus until reaching the maximum STR condition. Hence the onset of an engagement might be supersonic, and involve BVR weapons, but a classical WVR engagement, as described above, would develop rapidly. Therefore the majority of nominally ‘supersonic’ combat aircraft would spend a very small proportion of their flight time in the supersonic regime. The exception to this was the class of specialised point-defence interceptors, for which the short-duration dash comprised the whole of the outbound portion of their flight profile. However the overriding design issue for these aircraft was to minimise drag at supersonic conditions, resulting in limited flexibility for alternative missions. With the advent of more capable BVR weapons, the emphasis in combat has changed. These weapons are characterised by ‘no-escape-zone’ criteria, which are determined by the relative levels of energy in the weapon and the target. The size of the ’no-escape-zone’ depends on the energy imparted to the weapon on launch, as well as the onboard fuel. There is also the assumption that an adversary will also be equipped with similarly capable weapons. Hence having increased speed in combat increases the effectiveness of the BVR weapon, while reducing that of the threat. The ability to sustain supersonic speed to launch, and then to be able to turn and run from an incoming threat, becomes a factor in survivability. This results in a drive towards sustained manoeuvre capability at higher Mach number. The increased emphasis on supersonic performance for BVR combat results in demand for reduced wave drag due to both volume and lift, sufficient control power to overcome the significant trim changes through the transonic drag rise, attention to trim drag at manoeuvre conditions and reduced supersonic fuel burn, to allow for extended flight time in this regime. The nature of the flow at supersonic cruise conditions is predominantly steady and attached, with multiple interacting oblique shock structures around the aircraft. Areas of separated or unsteady flow are highly localised, and are mostly the consequence of bow shocks, with local pockets of subsonic flow behind them. The high energy of these flows means that where unsteadiness occurs, it can produce high frequency and amplitude pressure disturbances. This is particularly true in the vicinity of the intake system and the afterbody, where the jet efflux can interact strongly with both the aircraft boundary layers and the afterbody shock systems. V. Configuration Dependencies In categorising air vehicle configurations, we have followed the lead and terminology developed by Küchemann9 to describe families of aircraft. The type of flow around aircraft is fundamentally related to their configuration, both in terms of how they generate lift and control power, and also by the consequences for their relative moments of inertia about specific axes. Küchemann’s analysis, based on linear aerodynamic theory and some simplified assumptions about the relative proportions of different drag sources over the desired range of operating conditions, led to the observation that, as the predominant sources of lift-dependent drag at subsonic speeds were determined by span and at supersonic speeds by slenderness, the optimum span-to-length ratio for families of aircraft with successively higher cruise Mach conditions would reduce. Hence the optimum aerodynamic configuration for very low Mach numbers would be an unswept high aspect ratio wing, while the optimum for very high Mach number would be a slender body. For combat aircraft which are intended to operate in the high subsonic regime, the former case is represented by swept wing designs, e.g. F-86 and Hunter. For high Mach conditions, slender designs, e.g. F- 106 and SR-71, predominate. However, between these two extremes, the majority of more modern designs are either hybrids, e.g. F-16 and F-22, featuring aspects of swept and slender configurations, or have non-slender wings e.g. Eurofighter Typhoon. 10 American Institute of Aeronautics and Astronautics
  11. 11. Aside from the aerodynamic implications of increasing slenderness with Mach number, this has significant implications for the stability and control of these aircraft. For an aircraft of constant length, the change in optimum configuration with increasing Mach number leads to a progressively smaller moment of inertia about the roll axis, along with a consequent lack of moment arm for any control effector operating in that sense. Conversely, for an aircraft of constant span, the change in optimum configuration with increasing Mach number leads to a progressively larger moment of inertia about the pitch axis, along with an increased moment arm for control effectors in the same sense. A. Unswept/swept wing designs Aircraft designed to operate efficiently from low speeds up to the transonic regime will typically maximise aerodynamic performance by seeking to minimise profile drag, through reducing the wetted area required to contain a given volume, and minimise vortex drag by maximising wing span within structural mass constraints. In the extreme case, this drives the evolution of the classical sailplane configuration for operation at low Mach, involving slender, near-cylindrical fuselages and unswept wings with aspect ratios of 20 or higher. At sufficiently low Reynolds number, further reductions in profile drag are attainable by pursuing natural laminar flow. The nature of the wing flow for very high aspect ratio configurations is quasi-2D, with limited cross-flow effects being observed, and combinations of aerofoil characteristics, including viscous effects, and classical lifting line theory being generally capable of accurate design and analysis. The fluid mechanics of these designs are largely dictated by management of transition from laminar to turbulent and the development of turbulent flow separation from the trailing edge. This class of configuration forms the basis for most current UAVs designed for long endurance missions at low Mach number. Any requirement for increasing range requires higher operating speeds (i.e. maximising ML/D, rather than L/D) and hence the optimum configuration needs to reflect the associated increases in both dynamic pressure and Mach number. Increasing dynamic pressure increases the dimensional structural loads associated with wing bending moment. The effects of compressibility with increasing Mach number drive optimum wing thickness/chord ratio down. The combination of these factors produces a lower optimum aspect ratio for a given cruise requirement. Thus for aircraft designed to operate at M=0.6 to 0.7, the classical unswept configuration remains an option, as does natural laminar flow, although the increase in Reynolds number associated with speed increase makes this more demanding. The reduction in optimum aspect ratio makes localised three dimensional effects near the wing tip more significant, while the reduction in thickness makes flow separation at the trailing edge less, and at the leading edge more, likely. For design Mach numbers in excess of 0.7, compressibility effects become sufficiently important that combinations of further reduced wing thickness and increased sweep are necessary for efficient operation. Reduced wing thickness imposes a structural mass penalty for higher aspect ratio aircraft, and hence, for aircraft which are not required to operate above their transonic drag rise, the tendency is towards a combination of moderate sweep and aspect ratio. The trades between unswept and swept configurations for transport aircraft are highly dependent on the propulsion system, with higher bypass ratios, or open rotors, being more efficient at lower Mach. Improving aerodynamic and structural materials technology for civil transport aircraft has tended to drive towards increased aspect ratio and reduced sweep for a given cruise Mach number over time, but the trades between wing thickness and sweep are highly dependent on the balance between transonic wave drag and vortex drag at a required operating Mach number. For swept wings operating at these conditions, quasi-2D aerodynamic design begins to reach its limitations, and fully three-dimensional effects become more prevalent, particularly with regard to the development of boundary layers and locally supersonic flows. Hence, although the design pressure distributions for a moderately swept wing may be reasonably similar to those of a 2D aerofoil modified by means of sweep theory, the local boundary layer may feature significant cross-flow effects, particularly where faced by an adverse streamwise pressure gradient towards the trailing edge. This is the primary design space for large, turbofan-powered, civil transport aircraft. For thinner wings of higher sweep, spanwise boundary layer growth becomes an issue for leading edge flow separation. Thinner wings tend to generate higher leading edge suction peaks at off-design conditions, resulting in strong localised adverse pressure gradients immediately downstream of the leading edge. When combined with spanwise growth of the boundary layer, including along the attachment line, this increases the potential for leading edge flow separation outboard on the wing. For fixed leading edges, the adverse pressure gradient associated with the leading edge suction peak has tended to be reduced or controlled by means of twist, or ‘washout’, to accommodate the local upwash distribution, and local leading edge camber. For more modern combat aircraft designs, the adverse pressure gradient is managed by means of leading edge flaps, with the range of conditions over which flow separation is avoided expanded by appropriate scheduling. 11 American Institute of Aeronautics and Astronautics
  12. 12. The development of local areas of supersonic flow results in the propagation of disturbances along Mach cones within these regions, producing complex interacting three-dimensional flows. Shock waves will also develop, with the ideal being vanishingly weak oblique Mach waves, but unswept supersonic-to-subsonic shocks will also occur, generating wave drag, while the associated adverse pressure gradients will interact with boundary layers to generate a variety of non-linear flow phenomena, such as shock-induced separation and the resulting onset of unsteady flow instability (buffet onset)10. B. Slender wing designs Slender configurations are primarily designed for operation at supersonic conditions, where minimisation of wave drag, due to both volume and lift, is a design imperative. Hence slenderness ratio, rather than aspect ratio, is the driving factor in aerodynamic performance at these conditions. Although a variety of definitions of slenderness have been suggested, the most appropriate for design purposes is that the configuration lies entirely within the Mach cone originating from its own apex. For these configurations, the streamwise pressure gradients are necessarily much weaker than those in the associated cross-flow. Thus, for design purposes, the appropriate 2D analogue to the three-dimensional design problem is actually a conical flow, rather than a streamwise aerofoil. Aircraft designed for operation at supersonic conditions generally feature very low levels of thickness, to minimise wave drag due to volume, with values of thickness/chord of 4-5% being typical for combat aircraft. As at high subsonic conditions, thin wings tend to generate high local leading edge suction peaks where the flow normal to the leading edge is subsonic. These will tend to provoke flow separation, with the vorticity of the resulting free shear layer rolling up into a compact, coherent structure with a well-defined core and limited loss of streamwise momentum. Hence conical and near-conical vortex flows are a common characteristic of slender wings at lifting conditions. The existence of a primary vortex system generates significant cross-flow pressure gradients between the vortex core and the wing surface, resulting in significant cross-flow shear near the wall. Depending on the angle of incidence and local Mach number, the pressure gradients can produce embedded cross-flow shocks or other means of provoking secondary or even tertiary cross-flow separations and further associated vortex structures. The development of vortex flow topologies also creates a non-linear increase in normal force over the attached flow case, although this is more usually exploited at lower flight Mach numbers, as the increase in drag associated with the development of vortex flow structures penalises high speed performance very significantly. For this reason, aerodynamic design for this class of wings has evolved to minimise the drag at high speed conditions by eliminating the leading edge separation at the desired operating condition, usually a supersonic cruise or manoeuvre point. The conical camber methods originally used in the 1950s were largely supplanted by fully 3D potential flow methods during the 1960s, while modern designs use scheduled leading edge flaps to ensure efficient operation over a wider range of conditions. The development of shocks, flow separations and wakes in the conical sense is instructive for these configurations, and cross-flow flow topologies are not atypical of those arising in the fully three-dimensional reality. The conical flow analogy breaks down, however, when streamwise gradients become significant, particularly at the apex and base of these configurations. At the apex, there will be streamwise acceleration of the flow downstream of the stagnation point, itself potentially downstream of a bow shock. Hence there may be significant local streamwise gradients influencing the boundary layer transition behaviour at the apex and thus downstream and inboard of any primary vortex system. The streamwise adverse pressure gradient associated with the base will also have an impact on the development of the quasi-conical vortex flow, as the effect of a vortex core passing through an oblique or normal shock may cause either forward movement of the shock or induce vortex breakdown. Hence for slender configurations at high speed, shock-vortex interactions will introduce non-linear effects, analogous to the effects of shock-wave boundary-layer interactions on swept wings at transonic conditions. C. Hybrid and non-slender wing designs As discussed above, combat aircraft are expected to operate over a wide range of conditions, hence a configuration highly optimised for a single condition might be significantly penalised in other portions of the flight envelope. Part of the art of configuration aerodynamic design is thus to find an optimal compromise over a range of operating conditions. Design balance is therefore a major issue for combat aircraft. In practice, this drives towards prioritisation of some operating conditions against others and the emergence of aircraft design features to improve performance at specific points in the flight envelope. The majority of current combat aircraft designs are thus hybrids between swept and slender aircraft, with features of both, while non-slender wings have emerged as a compromise for aircraft where the need for high sweep must be balanced against increased aspect ratio for efficient subsonic cruise and manoeuvre. 12 American Institute of Aeronautics and Astronautics
  13. 13. While slender wing configurations offer reduced wave drag for supersonic operations, and the prospect of normal force increments due to vortex flow at lower Mach and high angle of incidence, their performance at transonic manoeuvre conditions is severely limited by their high vortex drag. Slender configurations will thus bleed energy and decelerate to low Mach number and altitude much more rapidly than a swept-wing equivalent, hence the evolution of configurations having a forebody strake or leading edge extension, combined with a thin wing of moderate leading edge sweep, typically lower than 40°, and moderate aspect ratio (typically 3-4). The resulting configurations tend to have much improved manoeuvre performance at high subsonic and transonic speeds, while gaining the benefit of the normal force increment at low speeds. The reduced slenderness of these configurations results in a supersonic wave drag penalty, which may be acceptable if the requirement for supersonic flight performance is limited. The flows over these configurations at subsonic manoeuvre conditions will be somewhat mixed, with local vortex flows inboard at lower speeds and outer wing flows featuring either leading edge or shock- induced flow separations at higher speeds. Combinations of both will arise over a broad range of Mach number, angle of incidence and altitude. For aircraft with a firmer requirement to operate at higher Mach number, such as BVR combat, higher sweep may be necessary to reduce wave drag to the necessary level. However, these aircraft will still require adequate turn performance at transonic conditions, thus aspect ratio remains an issue. To some extent, the increase in lift- dependent drag at transonic condition can be offset by increased thrust, but this then drives to a higher propulsion system mass, and reduced payload and fuel fraction. A configuration with a non-slender wing design is thus a system-level compromise. The nature of the flow over thin, non-slender wings is that they are subject to leading edge separations which are less predictable and coherently structured than those for slender wings. The leading edge flow separations may be either vortical or of the swept bubble-type, reflecting the greater loss of streamwise momentum associated with the non-slender leading edge flow separation. Growth of the boundary layer along the leading edge, including attachment-line transition, re-laminarization and crossflow transition phenomena, influences both the location and nature of the separation. In addition, use of scheduled leading edge flaps will influence the nature of the separation by changing the associated pressure gradients at both the leading edge and the flap knuckle. Hence the non-slender wing design incorporates all the most difficult features of the swept wing design, while introducing some flow features more usually observed on slender shapes. VI. Basis for the Taxonomy of Flight Mechanics Phenomena We are primarily concerned with addressing known risk areas, through a detailed understanding of flow physics, as part of an aircraft design process, rather than retrospective fixes for existing problems. In theory, if we are able to predict a specific fluid mechanics phenomenon using CFD, we should, in turn, be able to predict the consequences for air vehicle flight mechanics. However, in practice, as with wind tunnel and flight test, the number of feasible test points is determined by the budget available, while the complexity of CFD modelling required to address specific problems is not known a priori. Engineering judgement, although not infallible, is valuable as a guide to managing levels of risk, particularly when it is based on hard empirical evidence. The proposed basis for risk assessment is therefore based on a combination of first-order analysis derived from classical linear theory and empirical evidence derived from analysis of flight and wind tunnel measurements. Identifying which combinations of factors produce the highest risk should allow prioritisation of resource for accurate prediction of causative factors. It is proposed, based on the factors described above, to break down the problems of flight mechanics phenomena into categories: A. Flight mechanics mode: 1) Longitudinal: a. Steady-state b. Divergent c. Oscillatory 2) Lateral: a. Steady-state b. Divergent c. Oscillatory B. Flight regime: 1) Subsonic 2) Transonic 13 American Institute of Aeronautics and Astronautics
  14. 14. 3) Supersonic C. Manoeuvre type: 1) Cruise 2) Steady-state manoeuvre 3) Transient manoeuvre D. Configuration type: 1) Unswept 2) Swept 3) Slender 4) Hybrid swept/slender 5) Non-slender When taken in combination, these should form a matrix for which it is possible to assess the level of design risks and opportunities for improvement associated with each element, and to pursue diagnosis and palliative measures for the causation accordingly. VII. Fluid Dynamics Phenomena Underlying Flight Mechanics Issues Based on the discussion in Sections IV, V and VI, we can identify several fluid-dynamics phenomena that contribute to the establishment of flow fields around practical flight geometries and to the flight-mechanics issues listed above: A. Boundary-layer transition Several boundary-layer and free shear-layer transition modes can be identified for the flow fields of interest. Methods for control or promotion of these transition modes have been identified in the literature. It is noted that flow control approaches can be different for each of the following instability and transition modes. 1) Attachment-line instability and transition, affecting swept wings or slender fuselage noses, at high angle of incidence 2) Attachment-line contamination with inboard turbulence 3) Relaminarization and cessation of relaminarization of turbulent attachment-line flow, allowing attachment- line transition to appear. 4) Crossflow linear and non-linear instability and transition on wings and fuselage with sufficient sweep or body angle of incidence 5) Streamwise or Tollmien-Schlichting (T-S) instability and transition for boundary-layer edge Mach numbers below approximately 3.0 6) Second (Mack) mode inviscid instability and transition for boundary-layer edge Mach numbers above approximately 3.0 7) Shear-layer instability, transition and reattachment in laminar separation bubbles 8) Taylor-Görtler instability and transition B. Flow separation (no shock waves) Flow separation typically constitutes the largest driver for local and global changes in the flow field that cause modified flight-mechanics behaviour of the flight vehicle. Flow separation phenomena are ultimately linked to boundary-layer separation. A possible taxonomy of flow separation phenomena is as follows: 1) Leading-edge or trailing-edge separation (switchover depending mostly on geometry of airfoil and wing sweep as well as Reynolds number) 2) Laminar or turbulent state of boundary-layer upon separation (with transition and possible relaminarization in the separated shear layer) 3) Smooth surface versus “sharp-edged” shear layer separation 4) 2D vs. 3D type separation; open vs. closed separation topologies (Vortex vs. Bubble) 5) Junction and secondary flow separations 6) Off-surface flow reversal (in wake flows over multi-element airfoils). Impingement of wake-like flows on downstream lifting surfaces 7) Separation in Periodic flow field, including hysteresis effects 14 American Institute of Aeronautics and Astronautics
  15. 15. 8) “Unsteady” and “quasi-steady” separation C. Shock-wave/boundary-layer interactions Shock-wave/boundary-layer interactions are of concern as they can result in flow separation. They are proposed to be reviewed separately in current context as these interactions are strongly related to geometry and Mach regime of the flight vehicle. Shock-wave/boundary-layer interaction phenomena may be categorized along following observed flow phenomena: 1) Laminar boundary and turbulent boundary layer approaching the shock wave 2) Smooth surface vs. discontinuous surface (e.g. transonic shock vs. corner shock at supersonic edge Mach number) 3) Interactions on swept wings and non-swept wings 4) Separation bubble near the foot of the shock is closed or open (or, local vs. global separation) 5) Instability of flow field (‘steady’ vs. ‘unsteady’ interaction) 6) Buffet onset (global instability of the turbulent flow field that forces the (flexible) wing and fuselage structure) D. Shock-wave/vortex interactions 1) On slender wings: Shock-wave/vortex interaction on swept wings at supersonic speeds can result in rapid change of the topology of the strong leading edge vortex system as crossflow Mach number or angle of incidence are changed. Taxonomy of possible shock-wave/leading-edge-vortex interaction has been developed and published in the literature. 2) On Unswept/Swept wings: On unswept/swept wings (See Section V), shock-waves can interact with streamwise vortices generated by upstream vortex generators or chines that are embedded in the flow field approaching the shock-wave. E. Vortex stability, bursting and interactions Large changes in streamwise vortical flow can have large effects on the flow field surrounding the vortex. If the vortex is embedded in a region of flow deceleration, the vortex can burst. Following are some categories of flow fields with embedded vortical flows: 1) LEX/Chine vortex bursting and resulting fin buffet on Hybrid and Slender Wings 2) Forebody vortices from slender bodies interacting with downstream wing or empennage. 3) Foreplane tip vortex over downstream wing/empennage 4) Vortex from nacelle chine/strake flowing over unswept/swept wing with highly deflected flap settings F. Mixed-flow regions: spanwise segmentation of attached and separated flows Often it is unavoidable or desirable to combine regions of attached and separated flow fields. Using geometric or flow control devices, the flow on wings can be segmented into regions of attached and separated flow. Often the presence of a strong vortex can allow suitable segmentation of wing flow. Some possible segmentation examples: 1) Strake/swept wing with strong vortex on inboard strake and unswept type flow further outboard 2) Küchemann type tip flow field, where a stable vortical flow is generated on the outboard aft-swept wing tip, while the flow further inboard may be separated 3) Spanwise discontinuities in leading-edge geometry to affect span loading and formation of local vortices to provide spanwise containment of separated flow (e.g. drooped outboard leading edge, leading-edge notches, fences etc.) G. Flow control Flow control by unsteady flow manipulation and by geometric devices generally transfers higher momentum from outside the boundary layer towards the wall region, where the flow is separated or about to separate. Different control strategies may be needed for smooth-surface vs. sharp-edged separation. Active flow control by applying a periodic forcing, e.g. by synthetic jets, interacts with instability modes of separated shear layers, resulting in generation of vorticity relative to the flow that acts to entrain higher-momentum air towards the wall. The literature also provides guidance on where there are specific effects of scale on these various flow interactions, based on historical experience. In particular, an extensive survey of Reynolds number effects11 in 15 American Institute of Aeronautics and Astronautics
  16. 16. aircraft and weapons flows has been collated through NATO. This includes the treatment of specific problems through flow control. VIII. Causality – Which Fluid Phenomena? A. Structure of the taxonomic matrix Having generated a matrix of relevant factors in Section VI, it is necessary to identify sources of evidence to help identify causative fluid-dynamics factors from those described in VII for each combination of attributes. Ideally, these should be based on rigorous analysis of appropriate wind tunnel or flight test data. It is notable that the matrix representing the taxonomic factors from section VI is four-dimensional, and hence not immediately amenable to representation on a page or screen. However, as with flows themselves, some light can be shed on the problem by considering two-dimensional elements of the matrix. Considering the interaction between axes produces some simplifying generalisations, from which it can be deduced that the matrix we want to populate may be sparse enough for some preliminary conclusions to be drawn.: 1. Interaction between flight regime and configuration type It is a broad generalisation, supported by Küchemann’s analysis, that as the maximum operating Mach number of an aircraft increases, the configuration will necessarily become more slender. The resulting configuration must still then be able to operate safely over the entirety of the Mach number range down to launch and recovery conditions. A two-dimensional slice through this element of our taxonomic matrix will thus broadly appear to be triangular, with the diagonal consisting of a set of configurations operating near their maximum design Mach number. 2. Interaction between flight mechanics mode and configuration type Again, as a broad generalisation, issues of flight control about a given axis are complicated by increasing the bandwidth required to control any instabilities or non-linearities, and requiring sufficient control power (or moment arm) for effectors. Hence, as configurations become more slender, their moments of inertia about the roll axis become smaller, and the moment arm available decreases. Although this implies that roll-related problems decline in importance for less slender aircraft, there remain the coupled interactions about the yaw axis. Changing slenderness has somewhat of a lesser effect on the moment of inertia about the yaw axis. Hence less-slender aircraft will still be more susceptible to the spiral mode, while, conversely, more-slender aircraft will be increasingly susceptible to Dutch roll or wing rock. Changes in slenderness also have some impact on the susceptibility of configurations to longitudinal modes, as described above. In practice, this translates into a smaller useable centre of gravity range and a narrower band of acceptable static margin for unswept configurations. A plane through this aspect of the taxonomic matrix would therefore identify both roll and pitch characteristics as triangular with respect to configuration, while a corresponding examination of the yaw characteristics would show the matrix to be relatively dense. 3. Interaction between flight regime and manoeuvre type As considered in section IV, there is, for combat aircraft at least, significant interaction between the flight regime and manoeuvre type. In broad terms, at high Mach number, the angular rates generated are limited by both available thrust and structural loads, with operation beyond the thrust limit resulting in deceleration to lower Mach. Conversely, at the lowest end of the Mach number range, attained turns below corner point Mach are entirely high- rate transient phenomena. However, in the core part of the manoeuvring envelope, at transonic conditions between the maximum STR condition and maximum ATR condition, combinations of steady-state and time-dependent manoeuvre can be expected. For combat aircraft, this is the most important portion of the flight envelope, and one in which performance shortfalls have serious consequences. For civil transport aircraft, the requirements for operating away from a small number of specific operating conditions are relatively few. In normal operation, only cruise, takeoff and recovery phases are significant, although loiter performance is frequently required when holding before recovery. With the exception of some specific aspects of launch and recovery phases, these are steady-state conditions. However, the onset of non-linearity constrains the operation of these aircraft more severely than their military counterparts. Typically these will occur with increasing lift coefficient at either low-speed high-lift or high speed cruise with increasing altitude. B. Sources of data The vast bulk of detailed information on aerodynamic performance and, for combat aircraft at least, handling characteristics, is outside of the public domain and is either proprietary, subject to national security concerns, or both. Much of the information available in the public domain, particularly for simple configurations, is not 16 American Institute of Aeronautics and Astronautics
  17. 17. necessarily representative of, and thus relevant to, designed aircraft. For this reason, it is not particularly desirable to include such data. Some useful compendia of appropriate data do exist, and provide a basis to populate the matrix considered above with supporting evidence for causative factors. Some are historical in nature, providing an unclassified summary of more sensitive information for aircraft no longer in service. In the context of identifying causative factors, experimental data for cases where flight mechanics problems have been diagnosed and treated successfully by means of aerodynamic redesign or application of flow control are particularly valuable. Among the most valuable primary sources are experimental reports from official sources (e.g. NACA, NASA and, in the UK, ARC and RAE), which have been declassified or are suitable for downgrading. As the number of different aircraft programmes reduces decade-by-decade, and programme timescales become longer, the diversity of live experience in individual organisations declines. Familiarity with a broader set of historical data becomes a more important factor in ensuring an appropriate appreciation of where risks are likely to arise, and how they might be addressed. Within individual organisations, there are equivalent sources of information, which will also be of value. It will almost certainly be the case that data for the latest and most sensitive configurations are likely to be the least widely available and accessible. However, giving such data more lasting value, by ensuring that lessons are learnt and appropriately recorded, should be part of any responsible development programme. 1. Compendia of data A relatively short list of valuable summaries is described and discussed below. Note that these are not in themselves primary sources, and that this list is very far from being exhaustive. In some cases, typically where there has not been sufficient instrumentation (usually pressure measurement) in flight or wind tunnel tests, the level of analysis extends to driving flight mechanics, but not the underlying flow physics. A. AGARD-AR-82 This NATO advisory report12 represents the state of the art, circa 1975, in understanding of fluid mechanics phenomena related to WVR combat at transonic conditions. It contains contributions from across NATO nations, including much analysis of real aircraft experience, with a strong emphasis on the development of unsteady flow phenomena and their effect on flying qualities when engaged in combat. The report contains an extensive list of primary source material, which are analysed as a whole, and from which conclusions are drawn. This is probably the single most important secondary source of information relevant to causative factors in fluid dynamics, including descriptions of successful flow control solutions. B. AGARD-AR-155A This NATO advisory report13 follows on from the earlier reference12 and describes the implications of steady and unsteady separated flows for specific classes of aircraft configurations. As with the earlier document, the emphasis is on WVR combat at high subsonic speeds, and flying qualities within the tracking and evasion stages of engagements are of primary concern. It provides some guidance for incorporation of empirical design rules at the early stages of design. These empiricisms are largely based on experimental testing, and offer scope for re- investigation using more modern analysis methods. C. Abrupt Wing Stall Program The NASA/USN/USAF Abrupt Wing Stall (AWS) program was initiated due to problems experienced during flight tests of the F/A-18E. Major elements of the programme were presented in three special invited sessions as part of the 41st AIAA Aerospace Sciences Meeting and Exhibit at Reno in 2003, and subsequently in two dedicated editions14, 15 of the Journal of Aircraft. Part of the program involved a historical review of available data, and a summary16 of these was presented in both the meeting and journals. The review covered some of the data previously published through AGARD, described above, but also some more modern data. The historical review is predominantly focussed on lateral/directional characteristics in the high subsonic/transonic regime, where target tracking for WVR combat demands precise flying qualities. The AWS program also followed on from the previous AGARD studies in making recommendations for use of specific experimental and CFD techniques to evaluate the susceptibility of new designs to wing drop or wing rock at the high subsonic Mach numbers associated with sustained manoeuvre conditions. The recommendations are based on a more detailed analysis of the underlying flow physics, including steady and unsteady viscous CFD analyses, than was the case in the earlier AGARD summaries. 2. Specific experimental test programmes in the open literature There are a significant number of primary sources of flight test, or more often, wind tunnel test, data which can be related to specific issues encountered in flight. As the wind tunnel environment offers more control over test conditions, and frequently allows more detailed instrumentation to be installed, some wind tunnel data sets are very useful as a means of studying the fluid mechanics associated with particular configurations and flight regimes. The 17 American Institute of Aeronautics and Astronautics
  18. 18. wind tunnel environment is also the primary means of developing and validating flow control fixes. These are too numerous to list here, but some salient examples are: A. AGARD-AR-303 Volumes I & II This is the largest and most recent open compendium of wind tunnel data sets for validation of CFD methods17, building on earlier data sets collated through AGARD. B. International Vortex Flow Experiments (IVFE and VFE2) Two distinct programmes17, 18, 19, established to generate experimental force and pressure data at subsonic Mach numbers for uncambered 65° delta wings for validation of CFD methods. Experimental data exists with varying Reynolds number and leading edge radius. C. NASA F-16XL Flight Test Program The F-16XL flight test program20 provided in-flight pressure and boundary-layer measurements on a cranked- arrow wing over a wide range of operating conditions, including transonic and supersonic cases. D. NASA F/A-18 HARV Flight Test program The F/A-18 High Alpha Research Vehicle21 provided a basis for the in-flight study of vortex flows on a hybrid straked/swept wing configuration at low speed. The vehicle was used to study unsteady vortex bursting and associated fin buffet. E. NASA F-106B Flight-Test Program The F-106B flight test program22, 23 investigated the in-flight flow development on a cambered 60° delta wing, including the application of leading edge vortex flaps over a range of flight conditions. F. X-31 Flight Test Program The X-31 flight test program looked at post-stall manoeuvre for canard-delta configurations, including the use of thrust vectoring for control beyond the lift limit. G. RAE HP.115 flight test programme An aircraft designed and built for the specific objective of studying wing rock on slender delta wings. The vehicle became a testbed for extensive study of stability and control characteristics24 and vortex flows25. IX. Discussion and Suggested Topics for Future Workshops Although there is a place for blind comparisons with experimental data in the demonstration and validation of CFD tools being employed for specific programmes, a system of quality assurance for CFD based on taxonomy offers scope for learning from experimental and empirical experience. Based on an initial consideration of the taxonomy and data sources discussed in Section VIII, it is possible to suggest some of the most pressing fluid- mechanics phenomena (selected from the list in Section VII) that might be suitable candidates for specific workshops bringing together experimental data and CFD predictions under the auspices of one or more AIAA TC’s. These suggestions are largely based on their significance for operational use of combat and civil transport aircraft configurations. Further development of these topics will allow recommendation of suitably available existing or new geometry and flow data sets for the validation of specific CFD prediction capabilities. A. Transonic buffet The progressive development of transonic buffet is an issue for military and civil aircraft operating in the high subsonic flight regime. Buffet margin is a performance, structural life and flying qualities issue for both classes of aircraft. Hence a reliable capability for predicting the development of the associated fluid phenomena has a high payoff. The underlying flow physics are related to shock-wave boundary-layer interaction with progressively stronger shocks. The response of the boundary layer to the adverse pressure gradient provides a feedback mechanism. This development takes a series of phases: 1) A weak shock with limited thickening of the boundary layer downstream of the foot of the shock. 2) Thickening of the boundary layer leading to forward movement of the shock. 3) Further increases in shock strength, producing a steady flow separation downstream of the shock. 4) Unsteady flow separations downstream of the shock produce significant fore-and-aft shock movement. There are a number of 2D and 3D experimental and computational data sets available demonstrating aspects of this behaviour28-31. From the perspective of experiments in Shock-Boundary layer interactions, new data is now emerging from a variety of sources, including both fundamental ‘building-block’ experiments conducted on simple geometries with detailed flow measurements and tests on representative configurations. These represent a spectrum of cases that are suitable for use within a structured set of workshops. A recent activity within the European Union Sixth Framework programme, on “Unsteady effects of shock wave induced separation”, or UFAST32 is approaching completion. This has generated a series of detailed experimental 18 American Institute of Aeronautics and Astronautics
  19. 19. data sets for unsteady shock-wave boundary-layer interaction with a variety of geometries, along with a parallel set of CFD results using methods of varying complexity. This represents a significant set of ‘building-block’ experiments for this class of flows, and one potentially available for use within AIAA prediction workshops. A “Common Research Model” has recently been defined by a NASA and Industry Working Group for use in next AIAA Drag-Prediction Workshop. The geometry and to-be-taken wind-tunnel data at various Reynolds numbers will be released in the public domain33. Fig. 3 provides a view of the transport-type aircraft. The planned data sets include measurements at higher angles of incidence (including buffet onset and pitch up), in addition to cruise drag data for future Drag-Prediction Workshops. These existing and planned data sets offer the scope for progressive refinement of the CFD capability to address the need for prediction of unsteady shock-wave boundary layer interaction for representative 3D shapes. Future cooperative efforts between AIAA Technical Committees (notably APA, Fluid Dynamics and Atmospheric Flight Mechanics) and experts in industry, academia and government will need to develop further suitable experimental and computational geometries to investigate parts of this flow phenomenon. Methods to control onset of transonic buffet can be considered after correlating baseline buffet onset. Figure. 3 Proposed Common-Research Model (from Ref. 33) B. Flow separation from rounded leading edges Although there has been a great deal of progress in the prediction of separated flows from sharp leading edges, particularly for slender delta configurations, there has been far less progress for the prediction of flow separation onset for wing leading edges with finite radius and camber. Flow separation from the leading edge can have a significant adverse impact on both lift-dependent (induced) drag, due to loss of leading edge thrust, and stability characteristics, because of the associated rapid change in flow topology and associated centroid of wing lift. The delay and control of leading edge flow separation is a major factor in maximising transonic STR performance for combat aircraft. The lack of a reliable CFD capability for predicting this class of flows is offset by the extensive use of semi-empirical relationships for estimating attained suction limits. These are based on the results of appropriate experimental studies, which suggest a variety of contributory factors, including: 1) Transition from laminar to turbulent, with various transition mechanisms being significant, depending on the details of the geometry involved. T-S waves predominate at low sweep; cross-flow instability and attachment line transition become factors as sweep is increased. Engineering methods to predict the presence and significance of these boundary-layer transition phenomena are currently available and have been coupled to steady-state Reynolds-Averaged Navier-Stokes solution methods. 2) Stability of transitional (i.e. non-equilibrium) boundary layers in the presence of adverse pressure gradients. The development of a smooth-surface separation from a rounded wing leading edge is characterized by the relatively compact scale in which a variety of interacting phenomena occur. Three-dimensionality is very significant, with the behaviour of the leading edge flows on non-slender wings in particular showing significant sensitivity to very small changes in onset flow conditions and surface shape. The small scale of the flow phenomena and the sensitivity of these to contamination and other changes in boundary conditions makes the generation of appropriate experimental data fairly specific to particular cases. This is thus an area where selection of an experimental basis for a predictive workshop is distinctly non-trivial. Detailed fluid-mechanics prediction methods (including Direct Numerical Simulation) are becoming available to compute the sensitivity of cross flow and T-S receptivity, instability growth and transition to surface roughness, free stream turbulence, vertical and acoustical disturbances. 19 American Institute of Aeronautics and Astronautics
  20. 20. C. Flow separation upstream of the trailing edge Although flow separation from the wing trailing edge usually has less impact on drag and stability characteristics of an aircraft than a separation from the wing leading edge, it will have an impact on trailing edge devices used for either high lift or as control effectors. Although lag-entrainment boundary layer methods can predict 2D or unswept trailing edge flow separations very accurately, these are far less effective for predicting the fully 3D separations that emerge when trailing edges become more highly swept or when there are significant crossflow pressure gradients. As with the related problem of the secondary flow separations that occur under the influence of slender vortical flow, the relevant fluid dynamics are driven by the cumulative loss of momentum under the influence of shear and pressure gradients which may not be localised. Accurate prediction of three-dimensional smooth-surface flow separations therefore requires the full history of the boundary layer momentum to be captured in 3D, from attachment, through transition and adverse pressure gradients. Where the boundary layer is subject to localised unsteady pressures, the impact might be seen elsewhere (e.g. in the case of an unsteady trailing edge separation driven by the combination of an upstream shock and a local adverse pressure gradient). Three-dimensional flow separations from smooth surfaces represent the majority of secondary flows on aircraft. Their causation is not always directly associated with a localised effect, and yet their onset may have either subtle or rapid deleterious effects, not least on drag, but also potentially influencing stability and available control power to handle pitch-up and tuck-under characteristics on swept wings. These secondary flows are typically candidates for retrospective treatment with passive flow control techniques, such as vortex generators. As such, improvements in their prediction results in lower risk and need of rework. Indeed, this particular issue has been at the root of some of the uncertainty seen in the DPW series. X. Summary and Proposed Next Steps A. Summary It is assumed that CFD methods will increasingly form the primary method for the design and analysis of future combat air vehicles, with a consequent heavier reliance on their fidelity for more demanding regimes typically covered hitherto by wind tunnel or flight test. This implies that CFD methods need to be capable of predicting accurately the fluid mechanics phenomena of interest at these operating conditions. The objective of this White Paper is to provoke discussion within the AIAA Fluid Dynamics, Applied Aerodynamics and Atmospheric Flight Mechanics Technical Committees as to recommended directions in fluid dynamics research. These should be focussed towards specific flow phenomena which have problematic consequences for the flying qualities of combat and transport aircraft. It is proposed that a taxonomic approach is used to identify and prioritise these phenomena, primarily on the basis of existing empirical evidence, and where appropriate, identify where additional experimental evidence is required. A set of categories is proposed, comprising flight mechanics modes, flight regimes, required level of manoeuvre and air vehicle configuration, under which to identify and prioritise causative fluid mechanics issues. It is proposed that this build on previous work, conducted on the basis of designed air vehicle configurations and top-level user requirements. It is anticipated that the prioritisation of phenomena will provide a basis on which to conduct collaborative workshops covering the prediction of specific fluid dynamics topics, leading to an understanding of the capabilities of available CFD techniques and best practice in their use within the context of overall air vehicle aerodynamics. Three suggested topics for possible future Workshops have been identified; each topic has significant implications on vehicle flight mechanics and has adequate fluid dynamics and aerodynamic complexity for further study. The authors of this paper recommend focussing on one or two of the topics suggested in Section IX, i.e., transonic buffet and leading-edge separation. B. Proposed next steps Based on the authors’ preliminary estimation of steps needed to define possible Work Shops, it is recommended to proceed within the AIAA initially as follows: 1. Start a new Working Group (WG) within the AIAA’s Applied Aerodynamics (APA) and Fluid Dynamics (FD) Technical Committees to coordinate the effort proposed in this White Paper. It is suggested that the basis for a transonic buffet prediction workshop is strongest in the immediate near term. It is also suggested that building block experiments to support a leading edge separation workshop be identified in parallel. 20 American Institute of Aeronautics and Astronautics