AIAA White Paper on Fluid Dynamics Challenges in Flight mechanics

American Institute of Aeronautics and Astronautics1

Taxonomy of Flight Mechanics Issues for Aircraft, andunderlying 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 theAIAA Applied Aerodynamics Technical Committee, with contributions from interestedparties inside and outside AIAA. The objective of this paper is to define a process by whichthe 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 broaderuse of CFD methods to the aerodynamic design of air vehicles, assessments of theiraerodynamic 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 CFDmethods to predict correctly the results of ‘building block’ experiments. This White Paperlists several significant fluid dynamics phenomena that are commended topics for future(both experimental and computational) workshops under the auspices of relevant AIAATechnical Committees. The paper proposes steps within the AIAA to arrive at suchworkshops.

* 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

47th AIAA Aerospace Sciences Meeting Including The New Horizons Forum and Aerospace Exposition5 - 8 January 2009, Orlando, Florida

AIAA 2009-744

Copyright © 2009 by QinetiQ Ltd. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.


ContentsContents 2Nomenclature 3I. Introduction 3II. Flying Qualities Requirements 4

A. Certification standards 4B. Design aspects of flying qualities 5

III. Flight Mechanics Phenomena (“Flight Mechanics 101”) 6A. Longitudinal characteristics 6B. Lateral characteristics 7C. General characteristics of non-linear stability modes 7

IV. Taxonomy of Dominant Flow Physics and Flight Regimes for Aircraft 8A. Low speed and subsonic flight 9B. Transonic flight 9C. Supersonic flight 10

V. Configuration Dependencies 10A. Unswept/swept wing designs 11B. Slender wing designs 12C. Hybrid and non-slender wing designs 12

VI. Basis for the Taxonomy of Flight Mechanics Phenomena 13A. Flight mechanics mode: 13B. Flight regime: 13C. Manoeuvre type: 14D. Configuration type: 14

VII. Fluid Dynamics Phenomena Underlying Flight Mechanics Issues 14A. Boundary-layer transition 14B. Flow separation (no shock waves) 14C. Shock-wave/boundary-layer interactions 15D. Shock-wave/vortex interactions 15E. Vortex stability, bursting and interactions 15F. Mixed-flow regions: spanwise segmentation of attached and separated flows 15G. Flow control 15

VIII. Causality – Which Fluid Phenomena? 16A. Structure of the taxonomic matrix 16B. Sources of data 16

IX. Discussion and Suggested Topics for Future Workshops 18A. Transonic buffet 18B. Flow separation from rounded leading edges 19C. Flow separation upstream of the trailing edge 20

X. Summary and Proposed Next Steps 20A. Summary 20B. Proposed next steps 20

Acknowledgements 21References 21


NomenclatureAPA = Applied AerodynamicsAFM = Atmospheric Flight MechanicsATR = Attained Turn RateAWS = Abrupt Wing StallBVR = Beyond Visual RangeD = DragDPW = Drag Prediction WorkshopEASA = European Aviation Safety AgencyFAA = Federal Airworthiness AdministrationFCS = Flight Control SystemFD = Fluid DynamicsL = LiftM = Mach numberSPO = Short Period OscillationSTR = Sustained Turn RateTC = Technical CommitteeUAV = Uninhabited Air VehicleWVR = Within Visual Rangeω = Turn rate

I. IntroductionN recent years, the AIAA Applied Aerodynamics (APA) Technical Committee (TC) has sponsored a successionof 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 ofindustrial CFD use, in identifying technical issues in CFD application which need to be addressed, as a relativebenchmarking activity for practitioners, and, not least, as a forum in which future research needs could be brought tothe attention of appropriate funding sources.

Based on the positive experience from the DPW series, the military aviation community within the APA TCconsidered the feasibility of a similar moment prediction workshop series. This did not come to fruition, for avariety of reasons, not least the availability of a suitable experimental data set upon which to base a CFD predictionworkshop. In addition, there were concerns, based on a number of prior activities in this field, including the DPWseries, that any failure to predict some aspects of flight mechanics with CFD might leave participants none the wiseras to the underlying causes.

Issues have arisen in the DPW series, where specific fluid dynamics problems, such as development of trailingedge separations or localised separation bubbles, have not been solved by progressive global refinement of themeshes used for CFD1. Thus it appears that for some of these problems, the devil is in the detail, and it is unclearwhether 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 inspace and time. Experience from the DPW series, based on relatively ‘clean’ wind tunnel models, points towards theuse of successively denser meshes to spatially resolve local fluid dynamics issues. The feasibility and affordabilityof this approach for more complex configurations, at flight Reynolds numbers, are therefore questionable withoutthe continued operation of Moore’s law, which suggests that, with the passage of time, more computing power forthe same price will cure all ills.

There are other challenges in the world of military aircraft. The relative maturity of CFD for these applications issomewhat less than in the world of civil transport design. The range of operating conditions, and configurations, forcombat aircraft in manoeuvring flight is much more diverse than for the civil aircraft cruise cases considered withinthe DPW series. It is also true that the range of configurations and conditions encountered in weapons flight is yetmore diverse still. Given the issues identified within the DPW series, and the increased scope for local problems oncomplex military aircraft configurations, the global mesh refinement approach might prove to be a costly way toproceed, with no guarantee of success. It was unlikely that a ‘one-size-fits-all’ approach would cover all areas of themilitary aircraft flight envelope in which it is important to predict flight stability characteristics, hence an alternativeapproach was considered. This is based on a building-block approach for specific fluid dynamics phenomena, to

I


determine the most appropriate level of modelling for each, through a series of workshops combining analysis ofexperimental data and evaluation of CFD methods.

The adoption of a building-block approach to maturing CFD also recognises that, in many cases, the underlyingflow 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 releasabilityissues, and thus are open to a broader range of participants from industry and academia. This is not the first time thata ‘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 advocatedover a period of many years, notably through the NATO Research and Technology Organisation2. This approach ismore 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, henceproviding 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 thosespecifically determined by their inertia properties, a consequence of the aerodynamic loads to which they aresubject. Through the simple application of Newtonian mechanics, any change in the stability of an aircraft, whetherprogressive 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 theaircraft flow field. Structural elasticity under changing aerodynamic load may contribute to the non-linearity. To beable to predict the onset of non-linear stability characteristics with CFD therefore requires that the method should beable to predict the specific fluid phenomenon that provides causation. It is thus proposed that the means by which toaddress prediction of non-linear flight stability characteristics with CFD is to focus effort on prediction of specificfluid mechanics phenomena relevant to vehicle configurations under consideration.

The objective is to identify through taxonomy, for particular flight mechanics problems, the causative fluiddynamics phenomena, then determine, through a programme of workshops, the most appropriate level of modellingto capture the flow physics adequately. These workshops should seek to separate out model deficiencies fromnumerical issues. Where successful, these workshops should disseminate best practice for predicting thesephenomena in terms of both the level of physical modelling and discretization of the problem. Where unsuccessfulin predicting a particular phenomenon, appropriate levels of interest should be stimulated.

II. Flying Qualities Requirements

A. Certification standardsFlying 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 designand, where necessary and appropriate, flow control, can all make a contribution to a successful outcome. We areconcerned 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 theformal certification requirements.

This is expressed in outline in the relevant UK Defence Standard3: “It is desirable that the specified flying qualitiesshould be achieved by good aerodynamic and mechanical design. However automatic devices may be used where an overallbenefit 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 aircraftshall be…resistant to departure from controlled flight, post-stall gyrations and spins. Adequate warning of approach to departureshall be provided. The aircraft shall exhibit no uncommanded motion which cannot be arrested promptly by simple application ofpilot control.”

These certification standards, as is appropriate, have a good deal of commonality in their treatment of flyingqualities. Both contain more detailed qualitative and quantitative requirements and draw on a common basis forrating flying qualities: the Cooper-Harper ratings scale5, as shown in Figure 1. The Cooper-Harper rating processwas 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 embodiedin a Cooper-Harper rating 1-2, that pilot compensation is not a factor in achieving desired performance, is directlyapplicable to uninhabited air vehicles (UAVs), where autonomous flight, without remote pilot intervention, is aspecific design objective.

The certification standards for military aircraft are written for a wide variety of classes of air vehicle, from lightaircraft through to large multi-engine transport aircraft and surveillance platforms, as well as aircraft required to


perform manoeuvres over a large variety of operating conditions. However, the certification of light aircraft typesand transports for non-military uses is governed by the appropriate civil authorities. Within the EU these are theresponsibility 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 flyingqualities, With regard to stall characteristics: “It must be possible to produce and to correct roll and yaw by unreversed useof aileron and rudder controls, up to the time the aeroplane is stalled. No abnormal nose-up pitching may occur. The longitudinalcontrol force must be positive up to and throughout the stall. In addition, it must be possible to promptly prevent stalling and torecover from a stall by normal use of the controls” (CS 25.203). Likewise, for stall warning: “Stall warning with sufficientmargin to prevent inadvertent stalling with the flaps and landing gear in any normal position must be clear and distinctive to thepilot in straight and turning flight” (CS 25.207). In the case of buffeting: “The aeroplane must be demonstrated in flight to befree 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 beconsistent and predictable; that control departure boundaries should be well-defined, approached with adequatelevels of warning and easily recoverable from. The primary risks of failure are therefore associated with abruptand/or unpredictable changes in stability, and the lack of control power to overcome these.

B. Design aspects of flying qualitiesAs 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, eitherin mechanical systems or software. However, where increased performance demands and complexities in theunderlying aircraft stability characteristics arise, modern flight control systems offer a means of overcomingdeficiencies 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 theair vehicle system perspective, we must be aware of the issues which make FCS design more difficult and risky:

Is adequateperformanceattainable

with a tolerablepilot workload?

Pilot decisions

Is itcontrollable?

Is itsatisfactory

withoutimprovement?

Improvementmandatory

Deficiencies requireimprovement

Deficiencies warrantimprovement

10. Major deficiencies: Control will be lost during someportion of required operation.

8. Major deficiencies: Considerable pilot compensation isrequired for control.

9. Major deficiencies: Intense pilot compensation isrequired to retain control.

7. Major deficiencies: Adequate performance notattainable with maximum tolerable pilot compensation.Controllability not in question.

6. Very objectionable but tolerable deficiencies:Adequate performance requires extensive pilotcompensation.

5. Moderately objectionable deficiencies: Adequateperformance requires considerable pilotcompensation.

4. Minor but annoying deficiencies: Desired performancerequires moderate pilot compensation.

3. Fair – some mildly unpleasant deficiencies: Minimalpilot compensation required for desired performance.

2. Good - negligible deficiencies: Pilot compensation nota factor for desired performance.

1. Excellent – highly desirable. Pilot compensation not afactor for desired performance.

No

No

No

Yes

Yes

Yes

Is adequateperformanceattainable

with a tolerablepilot workload?

Pilot decisions

Is itcontrollable?

Is itsatisfactory

withoutimprovement?

Improvementmandatory

Deficiencies requireimprovement

Deficiencies warrantimprovement

10. Major deficiencies: Control will be lost during someportion of required operation.

8. Major deficiencies: Considerable pilot compensation isrequired for control.

9. Major deficiencies: Intense pilot compensation isrequired to retain control.

7. Major deficiencies: Adequate performance notattainable with maximum tolerable pilot compensation.Controllability not in question.

6. Very objectionable but tolerable deficiencies:Adequate performance requires extensive pilotcompensation.

5. Moderately objectionable deficiencies: Adequateperformance requires considerable pilotcompensation.

4. Minor but annoying deficiencies: Desired performancerequires moderate pilot compensation.

3. Fair – some mildly unpleasant deficiencies: Minimalpilot compensation required for desired performance.

2. Good - negligible deficiencies: Pilot compensation nota factor for desired performance.

1. Excellent – highly desirable. Pilot compensation not afactor for desired performance.

No

No

No

Yes

Yes

Yes

Figure 1 - Cooper-Harper Handling Qualities scale (from NASA TN-D 5153)


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 forinitial configuration layout, rather than the details of aerodynamic design and flow development. The first andsecond 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 inaerodynamic loads, although inertia loads have historically been responsible for control departure of specific classesof air vehicle7. Non-linearity of control response is frequently associated with variations in the aerodynamic loadassociated with the associated effectors, particularly where these are close-coupled and subject to significantinterference 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 massproperties, rather than through aerodynamic fixes.

It is apparent that, at the initial stages of a project, the primary influence of aerodynamic design on flight controlrisks 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 implicitdesign objective, the principles of consistency, predictability and avoidance of non-linear aerodynamic phenomenashould be borne in mind. However, as with flight control system design, where system performance demands requirecompromises, it is not always possible to achieve desired flying qualities through aerodynamic design alone. Anaerodynamic design based on consideration of the required flying qualities, however, remains the basis for allsubsequent palliatives and improvements derived from flight control systems and any flow control fixes. The betterthe 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 weare seeking to address.

Note that, given sufficient control power and actuation bandwidth, an unstable, yet linear, system is notparticularly problematic. Difficulties are more likely to arise when the underlying assumptions of linearity breakdown. In some cases, the non-linear stability modes which arise may be seen as variations of the linear stabilitymodes, particularly if periodic aerodynamic forcing occurs close to the frequencies associated with oscillatorymodes. However, there are any number of purely non-linear modes which can occur as the flight envelope requiredof 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 conditionsthen becomes a design issue.

A. Longitudinal characteristicsThe 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 longitudinalquartic are two complex conjugate pairs, representing two distinct oscillatory modes in pitch and heave. These areconventionally 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 highfrequency, while the Phugoid mode is a lightly damped oscillation of much lower frequency. However, theperformance benefits, in terms of reduced trim drag and increased agility, associated with reduced static marginmake it likely that both SPO and Phugoid modes can be unstable in their open-loop form, requiring automaticstabilization through feedback control. Meeting the overall system requirements then requires more effort in thedesign and flight clearance of appropriate control laws.


Moving on from the linear stability modes, there are a number of commonly experienced non-linear modes inpitch. 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 loadsobtained from wind-tunnel tests, while the oscillatory versions of these are generally the consequence of theassociated variation of static margin feeding through into variants of the existing linear modes. Large-amplitudepitch oscillations frequently develop hysteresis loops, further complicating their representation. Hence, moreaccurate, non-linear representations of the longitudinal modes should incorporate higher-order modelling than thefirst order linear stability derivatives to capture this behaviour adequately.

Uncontrollable pitch-up can and does have significant consequences for the useable flight envelope of anaircraft, particularly when at low speed and high lift. There are very many examples of aircraft subject to thisparticular problem, and, in practice, the usual solution is to limit the useable angle of incidence. The underlyingcauses 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 andnegative angles of incidence.

Both pitch-up and tuck-under are associated with a forward movement of the aerodynamic centre, and hencedecreased static margin. There is also the specific mode of Mach tuck, associated with the rearward movement of theaerodynamic centre with Mach number. This is of particular note at or near the transonic drag rise, resulting in asignificant increase in control power requirements and trim drag to achieve a given manoeuvre. In effect, thisbecomes a subtle limit on the useable angle of incidence at transonic and supersonic conditions, as the control powerrequired to trim increases beyond that available.

B. Lateral characteristicsBy 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 casestends to be sufficiently slow to be easily controllable by manual means. The roll subsidence mode is, likewise, aslow convergence or divergence in roll. Like the spiral, this is a weak mode in most practical cases. The Dutch rollmode is more significant than the two others, being an oscillatory coupling between motions about the roll and yawaxes. Notably, the stability of the Dutch roll and spiral modes interacts. In general, increases in the rolling momentdue to sideslip tend to promote instability in the Dutch roll mode, while having the opposite effect on the spiralmode. The converse is true of the yawing moment due to sideslip. As the Dutch roll mode is of relatively greaterimportance 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 berelated to the underlying linear modes, although at more extreme flight conditions, purely non-linear and coupledmodes 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. Bothof 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 canbe impacted by non-linearity. In particular, the oscillatory wing rock mode evolves from the Dutch roll mode whenaerodynamic damping terms fall to zero or become negative, with peak response being shown for specific, narrowfrequency bands.

C. General characteristics of non-linear stability modesIn 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-linearcharacteristics. The risk to a successful engineering solution providing acceptable flying qualities is determined bythe extent of the non-linearity and the rapidity of onset. Hence, at this point, it is desirable to address the fivespecific 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.


This list is by no means comprehensive. The classical linearised stability quartics are based on a number ofsimplifications and assumptions which do not hold across the full extent of the flight envelope, and hence the fullnon-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 AircraftAs 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 canrange from very high lift coefficients for launch and recovery at low Mach number through to very low liftcoefficients at the opposite end of the Mach number envelope and low altitude. Manoeuvre conditions for moderncombat aircraft are now much broader than the traditional Within-Visual-Range (WVR) transonic sustainedmanoeuvre 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 differentdominant flow physics, while the non-linear stability modes discussed above can be driven by completely differentflow mechanisms at different flight conditions. It is thus worth considering the varying nature of the flow fields atdifferent operating conditions, and the implications for developing taxonomy in which operating conditions are afactor. Figure 2, from ref. 8, shows a plot of turn rate, ω, vs. Mach number for three supersonic combat aircraft withrepresentative combat mass at the tropopause. Note that the peak sustained turn rate (STR) is usually, but notalways, at high subsonic Mach number, depending on the design requirements for a particular aircraft. The STR foreach of the aircraft shown in fig. 2 is thrust-limited at this altitude and significantly below the structural limit, whichmight apply at lower altitude. Peak STR is thus determined by the combination of propulsion system and aircraftdrag. Increasing thrust/weight ratio or reducing drag will improve the STR of the aircraft.

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 designdrivers are thus different, although many issues associated with handling qualities are common. It is perfectlyacceptable (and desirable) to perform manoeuvres with a combat aircraft which would not be the case with a civiltransport aircraft.

Figure 2 – Sustained Turn Rate vs. Mach number for a set of combat aircraft at the tropopause (from ref. 8)


A. Low speed and subsonic flightAt 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 bylow-speed requirements. Operations below minimum drag speed also offer the prospect of speed instability, andhence throttle response becomes an issue. For both civil transport and combat aircraft, low speed operations revolvearound launch and recovery, and for combat aircraft, to a more limited extent, low speed manoeuvre. High lift is afactor 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, bymeans of a runway, occasionally with assisted takeoff and recovery. Takeoff requires acceleration to rotation speedat maximum mass within a specified distance and hence drag is an important factor. It is therefore desirable that theflow over the aircraft at these conditions is close to fully attached, to minimise drag during acceleration to rotationspeed, and subsequently, on climb-out. Thus take-off is generally the least demanding low speed flight condition interms of flow complexity, although it is frequently the case that control effectors in pitch are sized by the need to liftthe nose wheel on rotation.

To recover an aircraft within the constraints of a required field length, or on the deck of an aircraft carrier, drivesminimum speed, and hence maximum lift coefficient. Issues of tail scrape and pilot visibility limit the useable angleof 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 landingspeeds there is a requirement to be able to sustain a trimmed angle of sideslip, as well as incidence. Drag is muchless 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 flightcondition 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 arecharacterised by two different flight regimes: lift-limited, where manoeuvre occurs up to maximum useable lift, andpost-stall manoeuvre, where angular rates are still being generated beyond maximum lift, using alternative controleffectors, such as vectored thrust. These manoeuvres tend to occur at Mach numbers below the corner point, atwhich attained turn rate (ATR) is a maximum. They are thus low-energy, transient manoeuvres, characterised byhigh angular rates, therefore the onset flow conditions are changing rapidly. The flows resulting at these conditionsare highly separated and unsteady. Pitch-down recovery from these conditions competes with the nose wheel liftrequirement to size the primary control effectors in pitch.

B. Transonic flightLike 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 (maximumL/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 toengage, followed by either WVR or BVR engagement, in which turns are made and conservation of energy is amajor factor.

In WVR combat, maximising turn rate, while minimising energy loss, drives towards aircraft operating at or neartheir peak STR for a given altitude. Potential energy can be traded for kinetic energy by losing altitude, withdisengagement becoming an issue as altitude, and hence Mach number, reduces. Excursions above the STR limitallow tighter turn radius, and thus potentially target acquisition and engagement, at the cost of energy bleed-off andsubsequent deceleration. Typical Mach numbers for maximum STR at the tropopause tend to be in the high subsonicregime. Openly published values8 of Mach number for peak STR at the tropopause for the F-16, MiG-21 and F-4 areapproximately 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 primaryair-superiority role. Drag minimisation at moderate to high lift coefficient is a design objective here, with the designlift coefficient at sustained manoeuvre being typically over 4 times that for cruise at the same Mach number andaltitude. The vast majority of the drag at these conditions is therefore lift-dependent. Hence, avoidance of flowseparation is desirable, but not always feasible.

At high subsonic manoeuvre cases, compressibility has a significant impact on the nature of the flow and on flowseparation mechanisms. Strong shock waves occur and, depending on the details of the configuration, become theprimary sources of both lift-dependent drag and flow separation as Mach number increases. Flows at transonicmanoeuvre conditions therefore tend to be separating to separated, with embedded shocks and consequent steady


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 theBreguet range equation, requires flight at conditions near or slightly into the transonic drag rise. However, therequirements for avoidance of buffet (measured at the pilot’s station in the cockpit) are more stringent than forcombat aircraft. This means that the onset of shock-induced flow separation, or trailing edge separation downstreamof a shock, is a condition to be avoided in normal flight envelope operations, including moderate manoeuvres, asdefined in airworthiness requirements. A civil transport aircraft operating at these conditions will generally tend tohave a wing loading somewhat higher than a combat aircraft operating at the same altitude and Mach number. Thecivil 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 flightUntil 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. Uponcommencing a manoeuvre, the aircraft would decelerate along the thrust-limited turn locus until reaching themaximum STR condition. Hence the onset of an engagement might be supersonic, and involve BVR weapons, but aclassical 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. Theexception to this was the class of specialised point-defence interceptors, for which the short-duration dash comprisedthe whole of the outbound portion of their flight profile. However the overriding design issue for these aircraft wasto 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 arecharacterised by ‘no-escape-zone’ criteria, which are determined by the relative levels of energy in the weapon andthe target. The size of the ’no-escape-zone’ depends on the energy imparted to the weapon on launch, as well as theonboard 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 thethreat. 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 Machnumber.

The increased emphasis on supersonic performance for BVR combat results in demand for reduced wave dragdue to both volume and lift, sufficient control power to overcome the significant trim changes through the transonicdrag rise, attention to trim drag at manoeuvre conditions and reduced supersonic fuel burn, to allow for extendedflight time in this regime. The nature of the flow at supersonic cruise conditions is predominantly steady andattached, with multiple interacting oblique shock structures around the aircraft. Areas of separated or unsteady floware highly localised, and are mostly the consequence of bow shocks, with local pockets of subsonic flow behindthem. The high energy of these flows means that where unsteadiness occurs, it can produce high frequency andamplitude 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 DependenciesIn 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, bothin terms of how they generate lift and control power, and also by the consequences for their relative moments ofinertia about specific axes. Küchemann’s analysis, based on linear aerodynamic theory and some simplifiedassumptions 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 byspan and at supersonic speeds by slenderness, the optimum span-to-length ratio for families of aircraft withsuccessively higher cruise Mach conditions would reduce. Hence the optimum aerodynamic configuration for verylow Mach numbers would be an unswept high aspect ratio wing, while the optimum for very high Mach numberwould be a slender body. For combat aircraft which are intended to operate in the high subsonic regime, the formercase 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 eitherhybrids, e.g. F-16 and F-22, featuring aspects of swept and slender configurations, or have non-slender wings e.g.Eurofighter Typhoon.


Aside from the aerodynamic implications of increasing slenderness with Mach number, this has significantimplications for the stability and control of these aircraft. For an aircraft of constant length, the change in optimumconfiguration 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 anaircraft of constant span, the change in optimum configuration with increasing Mach number leads to aprogressively larger moment of inertia about the pitch axis, along with an increased moment arm for controleffectors in the same sense.

A. Unswept/swept wing designsAircraft 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 containa given volume, and minimise vortex drag by maximising wing span within structural mass constraints. In theextreme case, this drives the evolution of the classical sailplane configuration for operation at low Mach, involvingslender, near-cylindrical fuselages and unswept wings with aspect ratios of 20 or higher. At sufficiently lowReynolds number, further reductions in profile drag are attainable by pursuing natural laminar flow. The nature ofthe 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 beinggenerally capable of accurate design and analysis. The fluid mechanics of these designs are largely dictated bymanagement of transition from laminar to turbulent and the development of turbulent flow separation from thetrailing edge. This class of configuration forms the basis for most current UAVs designed for long endurancemissions 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 Machnumber. Increasing dynamic pressure increases the dimensional structural loads associated with wing bendingmoment. The effects of compressibility with increasing Mach number drive optimum wing thickness/chord ratiodown. The combination of these factors produces a lower optimum aspect ratio for a given cruise requirement. Thusfor aircraft designed to operate at M=0.6 to 0.7, the classical unswept configuration remains an option, as doesnatural laminar flow, although the increase in Reynolds number associated with speed increase makes this moredemanding. The reduction in optimum aspect ratio makes localised three dimensional effects near the wing tip moresignificant, while the reduction in thickness makes flow separation at the trailing edge less, and at the leading edgemore, likely.

For design Mach numbers in excess of 0.7, compressibility effects become sufficiently important thatcombinations of further reduced wing thickness and increased sweep are necessary for efficient operation. Reducedwing thickness imposes a structural mass penalty for higher aspect ratio aircraft, and hence, for aircraft which arenot required to operate above their transonic drag rise, the tendency is towards a combination of moderate sweep andaspect ratio. The trades between unswept and swept configurations for transport aircraft are highly dependent on thepropulsion system, with higher bypass ratios, or open rotors, being more efficient at lower Mach. Improvingaerodynamic and structural materials technology for civil transport aircraft has tended to drive towards increasedaspect ratio and reduced sweep for a given cruise Mach number over time, but the trades between wing thicknessand sweep are highly dependent on the balance between transonic wave drag and vortex drag at a required operatingMach number.

For swept wings operating at these conditions, quasi-2D aerodynamic design begins to reach its limitations, andfully three-dimensional effects become more prevalent, particularly with regard to the development of boundarylayers and locally supersonic flows. Hence, although the design pressure distributions for a moderately swept wingmay be reasonably similar to those of a 2D aerofoil modified by means of sweep theory, the local boundary layermay feature significant cross-flow effects, particularly where faced by an adverse streamwise pressure gradienttowards 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 flowseparation. Thinner wings tend to generate higher leading edge suction peaks at off-design conditions, resulting instrong localised adverse pressure gradients immediately downstream of the leading edge. When combined withspanwise growth of the boundary layer, including along the attachment line, this increases the potential for leadingedge flow separation outboard on the wing. For fixed leading edges, the adverse pressure gradient associated withthe leading edge suction peak has tended to be reduced or controlled by means of twist, or ‘washout’, toaccommodate the local upwash distribution, and local leading edge camber. For more modern combat aircraftdesigns, the adverse pressure gradient is managed by means of leading edge flaps, with the range of conditions overwhich flow separation is avoided expanded by appropriate scheduling.


The development of local areas of supersonic flow results in the propagation of disturbances along Mach coneswithin these regions, producing complex interacting three-dimensional flows. Shock waves will also develop, withthe 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 generatea variety of non-linear flow phenomena, such as shock-induced separation and the resulting onset of unsteady flowinstability (buffet onset)10.

B. Slender wing designsSlender 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, isthe driving factor in aerodynamic performance at these conditions. Although a variety of definitions of slendernesshave been suggested, the most appropriate for design purposes is that the configuration lies entirely within the Machcone originating from its own apex. For these configurations, the streamwise pressure gradients are necessarilymuch weaker than those in the associated cross-flow. Thus, for design purposes, the appropriate 2D analogue to thethree-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, tominimise wave drag due to volume, with values of thickness/chord of 4-5% being typical for combat aircraft. As athigh subsonic conditions, thin wings tend to generate high local leading edge suction peaks where the flow normalto the leading edge is subsonic. These will tend to provoke flow separation, with the vorticity of the resulting freeshear layer rolling up into a compact, coherent structure with a well-defined core and limited loss of streamwisemomentum. Hence conical and near-conical vortex flows are a common characteristic of slender wings at liftingconditions. The existence of a primary vortex system generates significant cross-flow pressure gradients between thevortex core and the wing surface, resulting in significant cross-flow shear near the wall. Depending on the angle ofincidence and local Mach number, the pressure gradients can produce embedded cross-flow shocks or other meansof provoking secondary or even tertiary cross-flow separations and further associated vortex structures. Thedevelopment of vortex flow topologies also creates a non-linear increase in normal force over the attached flowcase, although this is more usually exploited at lower flight Mach numbers, as the increase in drag associated withthe 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 eliminatingthe leading edge separation at the desired operating condition, usually a supersonic cruise or manoeuvre point. Theconical camber methods originally used in the 1950s were largely supplanted by fully 3D potential flow methodsduring the 1960s, while modern designs use scheduled leading edge flaps to ensure efficient operation over a widerrange of conditions.

The development of shocks, flow separations and wakes in the conical sense is instructive for theseconfigurations, 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 theapex and base of these configurations. At the apex, there will be streamwise acceleration of the flow downstream ofthe stagnation point, itself potentially downstream of a bow shock. Hence there may be significant local streamwisegradients influencing the boundary layer transition behaviour at the apex and thus downstream and inboard of anyprimary vortex system. The streamwise adverse pressure gradient associated with the base will also have an impacton the development of the quasi-conical vortex flow, as the effect of a vortex core passing through an oblique ornormal shock may cause either forward movement of the shock or induce vortex breakdown. Hence for slenderconfigurations at high speed, shock-vortex interactions will introduce non-linear effects, analogous to the effects ofshock-wave boundary-layer interactions on swept wings at transonic conditions.

C. Hybrid and non-slender wing designsAs 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 flightenvelope. Part of the art of configuration aerodynamic design is thus to find an optimal compromise over a range ofoperating conditions. Design balance is therefore a major issue for combat aircraft. In practice, this drives towardsprioritisation of some operating conditions against others and the emergence of aircraft design features to improveperformance at specific points in the flight envelope. The majority of current combat aircraft designs are thushybrids between swept and slender aircraft, with features of both, while non-slender wings have emerged as acompromise for aircraft where the need for high sweep must be balanced against increased aspect ratio for efficientsubsonic cruise and manoeuvre.


While slender wing configurations offer reduced wave drag for supersonic operations, and the prospect ofnormal force increments due to vortex flow at lower Mach and high angle of incidence, their performance attransonic manoeuvre conditions is severely limited by their high vortex drag. Slender configurations will thus bleedenergy and decelerate to low Mach number and altitude much more rapidly than a swept-wing equivalent, hence theevolution of configurations having a forebody strake or leading edge extension, combined with a thin wing ofmoderate leading edge sweep, typically lower than 40°, and moderate aspect ratio (typically 3-4). The resultingconfigurations tend to have much improved manoeuvre performance at high subsonic and transonic speeds, whilegaining the benefit of the normal force increment at low speeds. The reduced slenderness of these configurationsresults in a supersonic wave drag penalty, which may be acceptable if the requirement for supersonic flightperformance is limited. The flows over these configurations at subsonic manoeuvre conditions will be somewhatmixed, 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 sweepmay be necessary to reduce wave drag to the necessary level. However, these aircraft will still require adequate turnperformance 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 propulsionsystem mass, and reduced payload and fuel fraction. A configuration with a non-slender wing design is thus asystem-level compromise. The nature of the flow over thin, non-slender wings is that they are subject to leadingedge separations which are less predictable and coherently structured than those for slender wings. The leading edgeflow separations may be either vortical or of the swept bubble-type, reflecting the greater loss of streamwisemomentum associated with the non-slender leading edge flow separation. Growth of the boundary layer along theleading edge, including attachment-line transition, re-laminarization and crossflow transition phenomena, influencesboth the location and nature of the separation. In addition, use of scheduled leading edge flaps will influence thenature 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, whileintroducing some flow features more usually observed on slender shapes.

VI. Basis for the Taxonomy of Flight Mechanics PhenomenaWe 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 topredict a specific fluid mechanics phenomenon using CFD, we should, in turn, be able to predict the consequencesfor air vehicle flight mechanics. However, in practice, as with wind tunnel and flight test, the number of feasible testpoints is determined by the budget available, while the complexity of CFD modelling required to address specificproblems is not known a priori. Engineering judgement, although not infallible, is valuable as a guide to managinglevels of risk, particularly when it is based on hard empirical evidence. The proposed basis for risk assessment istherefore based on a combination of first-order analysis derived from classical linear theory and empirical evidencederived from analysis of flight and wind tunnel measurements. Identifying which combinations of factors producethe 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 phenomenainto categories:

A. Flight mechanics mode:1) Longitudinal:

a. Steady-stateb. Divergentc. Oscillatory

2) Lateral:a. Steady-stateb. Divergentc. Oscillatory

B. Flight regime:1) Subsonic2) Transonic


3) Supersonic

C. Manoeuvre type:1) Cruise2) Steady-state manoeuvre3) Transient manoeuvre

D. Configuration type:1) Unswept2) Swept3) Slender4) Hybrid swept/slender5) Non-slender

When taken in combination, these should form a matrix for which it is possible to assess the level of design risksand opportunities for improvement associated with each element, and to pursue diagnosis and palliative measuresfor the causation accordingly.

VII. Fluid Dynamics Phenomena Underlying Flight Mechanics IssuesBased 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 issueslisted above:

A. Boundary-layer transitionSeveral 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 thatflow 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 ofincidence

2) Attachment-line contamination with inboard turbulence3) 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 incidence5) Streamwise or Tollmien-Schlichting (T-S) instability and transition for boundary-layer edge Mach numbers

below approximately 3.06) Second (Mack) mode inviscid instability and transition for boundary-layer edge Mach numbers above

approximately 3.07) Shear-layer instability, transition and reattachment in laminar separation bubbles8) 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 toboundary-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 wingsweep as well as Reynolds number)

2) Laminar or turbulent state of boundary-layer upon separation (with transition and possible relaminarizationin the separated shear layer)

3) Smooth surface versus “sharp-edged” shear layer separation4) 2D vs. 3D type separation; open vs. closed separation topologies (Vortex vs. Bubble)5) Junction and secondary flow separations6) Off-surface flow reversal (in wake flows over multi-element airfoils). Impingement of wake-like flows on

downstream lifting surfaces7) Separation in Periodic flow field, including hysteresis effects


8) “Unsteady” and “quasi-steady” separation

C. Shock-wave/boundary-layer interactionsShock-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 regimeof the flight vehicle. Shock-wave/boundary-layer interaction phenomena may be categorized along followingobserved flow phenomena:

1) Laminar boundary and turbulent boundary layer approaching the shock wave2) 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 wings4) 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 interactions1) On slender wings:

Shock-wave/vortex interaction on swept wings at supersonic speeds can result in rapid change of thetopology of the strong leading edge vortex system as crossflow Mach number or angle of incidence arechanged. Taxonomy of possible shock-wave/leading-edge-vortex interaction has been developed andpublished 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 interactionsLarge 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 flowfields with embedded vortical flows:

1) LEX/Chine vortex bursting and resulting fin buffet on Hybrid and Slender Wings2) Forebody vortices from slender bodies interacting with downstream wing or empennage.3) Foreplane tip vortex over downstream wing/empennage4) 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 flowsOften 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 thepresence 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 outboard2) 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 separated3) 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-edgenotches, fences etc.)

G. Flow controlFlow 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. Differentcontrol strategies may be needed for smooth-surface vs. sharp-edged separation. Active flow control by applying aperiodic forcing, e.g. by synthetic jets, interacts with instability modes of separated shear layers, resulting ingeneration 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 flowinteractions, based on historical experience. In particular, an extensive survey of Reynolds number effects11 in


aircraft and weapons flows has been collated through NATO. This includes the treatment of specific problemsthrough flow control.

VIII. Causality – Which Fluid Phenomena?

A. Structure of the taxonomic matrixHaving 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 hencenot immediately amenable to representation on a page or screen. However, as with flows themselves, some light canbe shed on the problem by considering two-dimensional elements of the matrix. Considering the interaction betweenaxes produces some simplifying generalisations, from which it can be deduced that the matrix we want to populatemay be sparse enough for some preliminary conclusions to be drawn.:

1. Interaction between flight regime and configuration typeIt 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 muststill then be able to operate safely over the entirety of the Mach number range down to launch and recoveryconditions. A two-dimensional slice through this element of our taxonomic matrix will thus broadly appear to betriangular, with the diagonal consisting of a set of configurations operating near their maximum design Machnumber.

2. Interaction between flight mechanics mode and configuration typeAgain, 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 momentarm) for effectors. Hence, as configurations become more slender, their moments of inertia about the roll axisbecome smaller, and the moment arm available decreases. Although this implies that roll-related problems decline inimportance for less slender aircraft, there remain the coupled interactions about the yaw axis. Changing slendernesshas somewhat of a lesser effect on the moment of inertia about the yaw axis. Hence less-slender aircraft will still bemore susceptible to the spiral mode, while, conversely, more-slender aircraft will be increasingly susceptible toDutch roll or wing rock.

Changes in slenderness also have some impact on the susceptibility of configurations to longitudinal modes, asdescribed above. In practice, this translates into a smaller useable centre of gravity range and a narrower band ofacceptable static margin for unswept configurations. A plane through this aspect of the taxonomic matrix wouldtherefore identify both roll and pitch characteristics as triangular with respect to configuration, while acorresponding examination of the yaw characteristics would show the matrix to be relatively dense.

3. Interaction between flight regime and manoeuvre typeAs 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 availablethrust 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 betweenthe maximum STR condition and maximum ATR condition, combinations of steady-state and time-dependentmanoeuvre can be expected. For combat aircraft, this is the most important portion of the flight envelope, and one inwhich performance shortfalls have serious consequences.

For civil transport aircraft, the requirements for operating away from a small number of specific operatingconditions are relatively few. In normal operation, only cruise, takeoff and recovery phases are significant, althoughloiter performance is frequently required when holding before recovery. With the exception of some specific aspectsof launch and recovery phases, these are steady-state conditions. However, the onset of non-linearity constrains theoperation of these aircraft more severely than their military counterparts. Typically these will occur with increasinglift coefficient at either low-speed high-lift or high speed cruise with increasing altitude.

B. Sources of dataThe 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, orboth. Much of the information available in the public domain, particularly for simple configurations, is not


necessarily representative of, and thus relevant to, designed aircraft. For this reason, it is not particularly desirable toinclude such data.

Some useful compendia of appropriate data do exist, and provide a basis to populate the matrix considered abovewith supporting evidence for causative factors. Some are historical in nature, providing an unclassified summary ofmore 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 bymeans 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, NASAand, in the UK, ARC and RAE), which have been declassified or are suitable for downgrading. As the number ofdifferent aircraft programmes reduces decade-by-decade, and programme timescales become longer, the diversity oflive experience in individual organisations declines. Familiarity with a broader set of historical data becomes a moreimportant factor in ensuring an appropriate appreciation of where risks are likely to arise, and how they might beaddressed. 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 leastwidely available and accessible. However, giving such data more lasting value, by ensuring that lessons are learntand appropriately recorded, should be part of any responsible development programme.

1. Compendia of dataA 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 therehas not been sufficient instrumentation (usually pressure measurement) in flight or wind tunnel tests, the level ofanalysis extends to driving flight mechanics, but not the underlying flow physics.

A. AGARD-AR-82This 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 flowphenomena and their effect on flying qualities when engaged in combat. The report contains an extensive list ofprimary source material, which are analysed as a whole, and from which conclusions are drawn. This is probably thesingle most important secondary source of information relevant to causative factors in fluid dynamics, includingdescriptions of successful flow control solutions.

B. AGARD-AR-155AThis 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 emphasisis on WVR combat at high subsonic speeds, and flying qualities within the tracking and evasion stages ofengagements are of primary concern. It provides some guidance for incorporation of empirical design rules at theearly 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 ProgramThe 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 partof the 41st AIAA Aerospace Sciences Meeting and Exhibit at Reno in 2003, and subsequently in two dedicatededitions14, 15 of the Journal of Aircraft. Part of the program involved a historical review of available data, and asummary16 of these was presented in both the meeting and journals. The review covered some of the data previouslypublished through AGARD, described above, but also some more modern data. The historical review ispredominantly focussed on lateral/directional characteristics in the high subsonic/transonic regime, where targettracking for WVR combat demands precise flying qualities.

The AWS program also followed on from the previous AGARD studies in making recommendations for use ofspecific experimental and CFD techniques to evaluate the susceptibility of new designs to wing drop or wing rock atthe high subsonic Mach numbers associated with sustained manoeuvre conditions. The recommendations are basedon 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 literatureThere 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 testconditions, and frequently allows more detailed instrumentation to be installed, some wind tunnel data sets are veryuseful as a means of studying the fluid mechanics associated with particular configurations and flight regimes. The


wind tunnel environment is also the primary means of developing and validating flow control fixes. These are toonumerous to list here, but some salient examples are:

A. AGARD-AR-303 Volumes I & IIThis 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 varyingReynolds number and leading edge radius.

C. NASA F-16XL Flight Test ProgramThe 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 programThe 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 andassociated fin buffet.

E. NASA F-106B Flight-Test ProgramThe 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 ProgramThe 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 programmeAn 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 WorkshopsAlthough 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 taxonomyoffers scope for learning from experimental and empirical experience. Based on an initial consideration of thetaxonomy 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 specificworkshops 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 aircraftconfigurations. Further development of these topics will allow recommendation of suitably available existing or newgeometry and flow data sets for the validation of specific CFD prediction capabilities.

A. Transonic buffetThe 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 ofaircraft. Hence a reliable capability for predicting the development of the associated fluid phenomena has a highpayoff. The underlying flow physics are related to shock-wave boundary-layer interaction with progressivelystronger shocks. The response of the boundary layer to the adverse pressure gradient provides a feedbackmechanism. 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 nowemerging from a variety of sources, including both fundamental ‘building-block’ experiments conducted on simplegeometries with detailed flow measurements and tests on representative configurations. These represent a spectrumof 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 waveinduced separation”, or UFAST32 is approaching completion. This has generated a series of detailed experimental


data sets for unsteady shock-wave boundary-layer interaction with a variety of geometries, along with a parallel setof 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 innext AIAA Drag-Prediction Workshop. The geometry and to-be-taken wind-tunnel data at various Reynoldsnumbers will be released in the public domain33. Fig. 3 provides a view of the transport-type aircraft. The planneddata sets include measurements at higher angles of incidence (including buffet onset and pitch up), in addition tocruise drag data for future Drag-Prediction Workshops. These existing and planned data sets offer the scope forprogressive refinement of the CFD capability to address the need for prediction of unsteady shock-wave boundarylayer interaction for representative 3D shapes.

Future cooperative efforts between AIAA Technical Committees (notably APA, Fluid Dynamics andAtmospheric Flight Mechanics) and experts in industry, academia and government will need to develop furthersuitable experimental and computational geometries to investigate parts of this flow phenomenon. Methods tocontrol 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 edgesAlthough 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 separationonset for wing leading edges with finite radius and camber. Flow separation from the leading edge can have asignificant adverse impact on both lift-dependent (induced) drag, due to loss of leading edge thrust, and stabilitycharacteristics, because of the associated rapid change in flow topology and associated centroid of wing lift. Thedelay and control of leading edge flow separation is a major factor in maximising transonic STR performance forcombat aircraft. The lack of a reliable CFD capability for predicting this class of flows is offset by the extensive useof semi-empirical relationships for estimating attained suction limits. These are based on the results of appropriateexperimental studies, which suggest a variety of contributory factors, including:

1) Transition from laminar to turbulent, with various transition mechanisms being significant, depending on thedetails of the geometry involved. T-S waves predominate at low sweep; cross-flow instability and attachmentline transition become factors as sweep is increased. Engineering methods to predict the presence andsignificance of these boundary-layer transition phenomena are currently available and have been coupled tosteady-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 verysignificant, with the behaviour of the leading edge flows on non-slender wings in particular showing significantsensitivity to very small changes in onset flow conditions and surface shape. The small scale of the flow phenomenaand the sensitivity of these to contamination and other changes in boundary conditions makes the generation ofappropriate experimental data fairly specific to particular cases. This is thus an area where selection of anexperimental 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-Sreceptivity, instability growth and transition to surface roughness, free stream turbulence, vertical and acousticaldisturbances.


C. Flow separation upstream of the trailing edgeAlthough 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 foreither high lift or as control effectors. Although lag-entrainment boundary layer methods can predict 2D or unswepttrailing edge flow separations very accurately, these are far less effective for predicting the fully 3D separations thatemerge 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 vorticalflow, the relevant fluid dynamics are driven by the cumulative loss of momentum under the influence of shear andpressure gradients which may not be localised. Accurate prediction of three-dimensional smooth-surface flowseparations therefore requires the full history of the boundary layer momentum to be captured in 3D, fromattachment, through transition and adverse pressure gradients. Where the boundary layer is subject to localisedunsteady pressures, the impact might be seen elsewhere (e.g. in the case of an unsteady trailing edge separationdriven 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 orrapid deleterious effects, not least on drag, but also potentially influencing stability and available control power tohandle pitch-up and tuck-under characteristics on swept wings. These secondary flows are typically candidates forretrospective treatment with passive flow control techniques, such as vortex generators. As such, improvements intheir prediction results in lower risk and need of rework. Indeed, this particular issue has been at the root of some ofthe uncertainty seen in the DPW series.

X. Summary and Proposed Next Steps

A. SummaryIt 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 typicallycovered hitherto by wind tunnel or flight test. This implies that CFD methods need to be capable of predictingaccurately 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, AppliedAerodynamics and Atmospheric Flight Mechanics Technical Committees as to recommended directions in fluiddynamics research. These should be focussed towards specific flow phenomena which have problematicconsequences for the flying qualities of combat and transport aircraft. It is proposed that a taxonomic approach isused to identify and prioritise these phenomena, primarily on the basis of existing empirical evidence, and whereappropriate, identify where additional experimental evidence is required.

A set of categories is proposed, comprising flight mechanics modes, flight regimes, required level of manoeuvreand air vehicle configuration, under which to identify and prioritise causative fluid mechanics issues. It is proposedthat this build on previous work, conducted on the basis of designed air vehicle configurations and top-level userrequirements. It is anticipated that the prioritisation of phenomena will provide a basis on which to conductcollaborative workshops covering the prediction of specific fluid dynamics topics, leading to an understanding of thecapabilities of available CFD techniques and best practice in their use within the context of overall air vehicleaerodynamics.

Three suggested topics for possible future Workshops have been identified; each topic has significantimplications on vehicle flight mechanics and has adequate fluid dynamics and aerodynamic complexity for furtherstudy. 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 stepsBased 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 FluidDynamics (FD) Technical Committees to coordinate the effort proposed in this White Paper. It is suggestedthat the basis for a transonic buffet prediction workshop is strongest in the immediate near term. It is alsosuggested that building block experiments to support a leading edge separation workshop be identified inparallel.


2. The WG identifies and invites experts (“greybeards”) who can address the overall topic of CFDprediction and can help focus the WG and the community on topics of main interest at future AIAAAPA/FD conferences, via dedicated sessions with invited speakers and a discussion session. The“greybeards” are expected to indicate state of the art, challenges, opportunities and previous, related, effortsand results on the topics of interest.3. The WG identifies and contacts organizations that have existing data sets in the public domain thatcould be used for CFD correlation.4. The WG proposes a set, or sets, of geometry for the community to analyze, or, in the absence ofsuitable existing geometry and wind-tunnel data, proposes a new geometry for CFD analysis. The WGadvocates testing of the geometry at suitable governmental and/or university wind-tunnel facilities.5. After agreement on suitability of a data set, with existing or expected experimental data, the WGproposes to the TCs to hold a series of related workshops.

It is the opinion of the present authors that the actual achievement of a workshop on one of the topics suggestedin Section IX is only feasible if the members of the proposed new Working Group, under the auspices of thespawning Technical Committees, coordinate pro-actively within the entire community (not just AIAA) to ensure thata workshop aligns with possible related efforts in other organizations (e.g., RTO, GARTEUR, KATnet, othernational and international programmes as well as university research).

The purpose of the current proposed effort is to advance the state of the art, and not to compete for limitedresources, particularly in the area of wind tunnel testing, that are available in our field of work. The present authorsfirmly believe that a first workshop can be held in 2010 or 2011 on one of the proposed topics. It is up to the appliedand fluid dynamics communities within the AIAA, together with the larger aerodynamics community to make ithappen. The relative success of, and interest in, the AIAA Drag Prediction Workshops indicates the benefits ofcoordinated and specific workshop formats to make progress in the state of the art.

AcknowledgementsThanks are due to current and past members of the AIAA Applied Aerodynamics and Fluid Dynamics Technical

Committees for their involvement in evolving this White Paper. Thanks are due to Bram Elsenaar, John Fulker,Darren Grove, Frank Lynch and Jeff Slotnick, for constructive criticism. Particular thanks are due to Paul Vijgen forhis large contribution to organising this effort.

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AIAA White Paper on Fluid Dynamics Challenges in Flight mechanics