ADVANCED DESIGN II SUMMARY 1 Advanced Aircraft Design II: Summary

I.INTRODUCTION Progress in fighters: • Turbojet and swept (1940s) • Autostabilisation (1950s) • Guided weapons (1950s) • Leaky turbojets 1960s) • Microprocessors (1970s) • Fly-by-wire and artificial stability (1970s) • Composites (1970s) • Stealth (1980s) • Supermanoeuvrability (1990s)

Requirements: • Lethality Fig. 1. Drag breakdown • Manoeuvrability • Handling qualities External stores have a large effect on the flight envelope • Radius of action (flight envelope shrinks with stores), mainly due to increased • Persistence drag and possibly aeroelasic/interference effects. • Resilience • Visibility Airfoil design • Stealth General fact: thinner means heavier wings. Classfication of jet fighters 2nd generation fighters had thin wings for high speed flight, but this caused separation at subsonic maneuvers • 1st generation (mid-1940s to mid 1950s) and buffet at low angels of attack. This resulted in a bat firing • 2nd generation (mid-1950s to early 1960s) platform. There was a need for thicker airfoils with good • 3rd generation (early 1960s to circa 1970) transonic characteristics. The answer was the supercritical • 4th generation (1970 to mid 1990s) airfoil. The good characteristics were a result of a rapid flow • 4.5th generation (1990s to present) expansion about the leading edge and isentropic recompression • 5th generation (2005 to present) through beneficial wave interaction. Combat aircraft types: • Reconnaissance – Strategic reconnaissance (U2, SR-71) – Tactical reconnaissance (derivative of fighter) • Ground attack • Interceptors • Air superiority

II.AIRFOILANDWINGPLANFORM vortex drag: Fig. 2. Supercritical airfoil • 75% of total drag during maneuvering Supercritical airfoil: • 50% of total drag during cruise • Increase the drag-rise Mach number for a given thickness • 5-10% of total drag in low altitude, high speed flight ratio and sweep. • Allow use of thicker wing for a given MD and sweep in Profile drag: order to improve available wing volume and either reduce • friction drag (30% during cruise) wing structure weight or increase the aspect ratio. • form drag • Reduce wing sweep for a given MD and thickness ratio, so improving lift and lift/drag ratio for take-off and ADVANCED AIRCRAFT DESIGN II SUMMARY 2

– Gust response – small wing, high wing loading • So what do we do? – Find the smallest wing that meets requirements – Opt for variable sweep

Wing tips: • Kuchemann tip (Harrier) Fig. 3. Wave interaction – Good transonic characteristics • Raked tip (F-15) – can be loaded higher than expected landing. – Reduced bending moments – Reduced buffet Conical drag: – Increased dutch roll dampinning • Improve off-design performance of supersonic fighters. • Straight tip (F-16) • Suppress leading-edge separation by increasing buffet CL and postponing drag break. – Allows launcher rail – Could improve L/D Aspect ratio: • Effect on Taper ratio: – Trailing-edge vortex drag (aka induced drag) – Lift-curve slope • Taper ratio in combination with moderate sweep: – Low supersonic drag • High aspect ratio required for: – Increased spiral stability (C ) through leading edge – Long endurance (∝ L/D) l,β sweep – Long range (∝ ML/D) – Effective trailing-edge flaps – Subsonic maneuvering (up to certain AoA) – Low AoA requirement at take-off and landing • Cross-wind handling problems at high AoA • Drawbacks of high aspect ratio – Increased rolling moment due to sideslip, Cl,β) – Weight penalty – Less control power due to swept trailing edge – Supersonic drag increased • Reduces root bending moment and thus wing weight – Sensitive to atmospheric upsets • Higher loaded outboard sections Wing twist: • Higher possibility of tip if combined with sweep: pitch-up and wing drop • To prevent tip stall • To adjust spanwise loading and achieve minimum drag at a certain condition • (sometimes also to adjust the pitching moment) Swept and delta wings • (at high g-maneuvers aeroelastic bending causes Benefits of wing sweep: aerodynamic twist, up to 10 deg) • Inventors: Adolf Busemann, Albert Betz, Hans Multhopp • Velocity compnent perpendicular to the wing: V cot cos Λ Wing size: • Sweep delays drag rise and reduces peak drag • Gross wing size: • At subsonic speeds sweep penalises L/D – Large effect on drag • Reduction of tuck-under effect – Crucial role for sizing the aircraft (weight and thus – Supersonic patch results in shift aft shift of aerody- cost) namic center – Snowball effect of wing size on airplane size – Result = nose down pitching moment (tuck) • What drives wing size? – More gradual shift on a – Field performance – low wing loading desired for • More gradual variation of lift coefficient across the tran- short fields sonic region – Subsonic cruise and loiter – medium wing loading • Extension of buffet boundaries desired – Lower overspeeds at given Mach number and CL – Sustained turn rate – low wing loading (high A for – Less strong shockwave terminating supersonic patch low C ) D,i – Postponement of separation at the foot of the shock – Instantaneous turn rate – low wing loading (high CL) – High supersonic dash – high wing loading • Reduction of gust reponse (good for high speed penetra- – Subsonic SEP – not directly affected by wing size tion) – Low altitude & high speed – small wing, high wing – Sweep reduces CL,α loading • For thin, low aspect ratio wings: higher CL,max ADVANCED AIRCRAFT DESIGN II SUMMARY 3

– Stable vortex separation induces vortex lift up to high – Drag less sensitive to Mach number AoA – Easier to obtain satisfactory cross-sectional area dis- – Stall might be more gradual tribution (no HT)

• With sweep back, wave drag becomes: • Gradual change of CL and CLα with M – Independent of span loading • Leading-edge vortex gives better stall behavior – Linearly dependent on (t/c)2 • Allows light wings with high bending and torsional – Minimized by spreading lift over large chord stiffness • Wing can stay out of Mach cone (if sweep angle is large – Thicker wings allow for more volume for fuel and than Mach angle) gear – Flutter and aileron reversal can be eliminated Penalties of sweepback: • Low wing loading allows for acceptable maneuvering and • Limits of theory: handling – valid for infinitely long skewed wings • Smaller wings do not require folding – Flow perpendicular to isobars • Large wing area available for external stores – Root and tip effects dominant on low aspect ratio wings Disadvantages of delta wings • Result: • Tailless deltas have high landing speeds and bad field performance – Delay in Mcrit overpredicted – In practice half of the expected amount in Mcrit is – Low-lift curve slope requires high AoA possible – Tail clearance limits AoA • Lift curve slope increases, until vortex breaks down at – Unable to trim out the nose-down pitching moment the trailing edge from flaps • Loss of leading edge suction leads to increased lift- • High lift-induced drag in subsonic conditions dependent drag – High thrust required • With increasing leading edge radius the vortex will appear – Trimmed lift loss at high AoA due to downloading at higher angles of attack trailing-edge controls • Structural problems • Low wing loading • Reduced effectiveness of high-lift devices – Although CLα is low, Lα is high due to low W/S → • Tip stalling (especially with combination of sweep and gust response large aspect ratio) – High wing loading would compromize manoeuvra- • Increased rolling due to sideslip bility • Increased drag due to lift • Supersonic manoeuvrability restricted • (reduction of lift-curve slope Clα → nose-high attitudes at landing → raised cockpits required for visibility) – Trailing-edge flight controls () are less effec- • Rolling moment due to side-slip is increased due to tive sweep. – Large absolute shift in a.c. (needs to be trimmed and • Reduction of wing controls and flaps may demand c.g. shift) – leading-edge sweep dominant for its effective- • Excessive Clβ at low speed ness – Large (leading-edge) sweep and high AoA disturbs – High flap sweep angles reduce ∆CLmax desired relation between lateral and directional sta- bility, Dutch roll becomes exaggerated, low wing and As way to counter some of the downsides of sweep is the yaw dampers required. reduce the trailing-edge sweep and make the root chord • Pitch damping reduced (if there is no horizontal tail) larger (additional benefit here is that this strengthens the rear – Risk of pitch induced oscillation and central torsion box). So a good way to enhance – Pitch dampers might have to be installed supersonic maneuvering is to have a low aspect ratio, large – In case of horizontal tail use a low-mounted ht to wing. avoid deep stall at high AoA

Delta wing Unstable delta does have some other possible advantages, • Alexander Lippisch (1931), Avro Vulcan (1947), see Fig. 4. Dassault Mirage I (1952) Compound sweep delta (F-16 XL) Benefits of delta wings: • Longer • Transsonic drag rise is more gradual and peak supersonic • Twice the wing area drag is reduced – allows for more hard points – Lift spread over broader chord (lower section cl) – increases friction drag ADVANCED AIRCRAFT DESIGN II SUMMARY 4

Fig. 4. Advantages of unstable delta

• Increased fineness ratio and wing fuselage blending • Wing optimized for low-level supersonic speeds • Trailing edge reflex included • 70 deg leading edge within the shock cone of the nose • 50 deg swept outboard wing with thin profile and sharp leading edge F-16XL flying qualities: • Lateral/directional stability is improved • External loads no not adversely affect flying qualities In modern air combat fighters delta wings are used because the high degree of leading-edge sweep promotes strong vortex formation at high AoA. It has low wave drag at supersonic Fig. 5. Effect of glove size speeds and the combination with a foreplane creates beneficial interference. Arguments for inboard pivot: Variable swept wings, used for: • Fully swept wing area and span are smaller • Long-range subsonic cruise or long-endurance loiter on • gives the highest effective aspect ratio in the unswept station condition • High-supersonic interception and transonic low-altitude • more aeroelastic relieving effect on pitch stiffness. strike • trim drag penalty is not particularly acute for combat • Operation from limited-length runways or aircraft carri- aircraft using full sweep only for supersonic dash or low- ers. altitude, high-speed penetration of limited duration Using a variable sweep wing can also be used for low- • Trim change can be hidden from pilot by pitch dampers altitude high-speed action (F-111), with the wings swept back • The complications of fairing and sealing a fixed apex the wing has a lower aspect ratio and lower CLα so it is less are avoided, allowing the use of full-span leading-edge sensitive to gust upsets. High sweep will also bring the a.c high-lift devices. more aft and thus increase the corrective effect of Cm (pitch stiffness). When the low-altitude high-speed dash has been Observations of sweep wing: completed the aircraft can then benefit form the (low speed) • Increasing sweep from 25 deg to 67.5 deg decreases the advantages of an unswept, high aspect ratio wing (good lift curve slope by 50 percent. This considerably lessens take-off and landing performance, more efficient subsonic the susceptibility to gusts. cruise and loiter, better subsonic sustained manoeuvring). • Even at subsonic Mach numbers the stability increases greatly with increasing wing sweep notwithstanding the Disadvantages of variable sweep wings: use of a fixed glove on the inboard wing. At supersonic • Excessive static stability at high sweep (small c.g. excur- speeds stability increases even more. sion, large n.p. excursion) although reduced by aeroelastic • Increasing wing sweep from 25 deg to 65 and 67.5 deg effects... increases the drag rise Mach number from M = 0.75 to • Large trim drag due to aft a.c. (induced drag of wing and close to M = 0.90. horizontal tail) • Large deflections required at high AoA Forward swept wings: • Hence: large down force of tail should be compensated Main problem is the combination of bending (aerodynamic by larger lift twist) and torsion (geometric twist) that occurs when a forward • Even more aft a.c. at transonic conditions reduces ma- swept wing is constructed using an isotropic material. By neuverability using an anisotropic material one can decouple the bending Possible solutions are a translating wing or to move the and torsion to obtain a wing that does not diverge. (it is pivot point outboard. also possible with isotropic material, but the structure would become quite heavy) ADVANCED AIRCRAFT DESIGN II SUMMARY 5

Advantages of forward sweep: – High-speed boundary layer discharges from slat • Roll control and damping more effective at high AoA training edge • Reduced dihedral effect at high AoA – Reduces adverse pressure gradient on slat • Boundary layer drifts inboard (at high AoA) • Fresh boundary layer effect – Higher load inboard section – A new boundary layer is formed on each new com- – Inboard stall could create pitch-up (when behind ponent c.g.). (Fences, inboard twist or limited area of aft • Characteristics: sweep can prevent this) – Possible increase in leading-edge camber – Foreplane () can produce downwash to keep – Possible increase in chord the flow attached – Small change in Cm • For same sock sweep less leading edge sweep required • Higher aspect ratio → higher CLα → higher CL at take- Typical modern fighters have low CLmax and low CLα, off and landing when restricted by tail strike or pilot they are driven by speed and weight requirements, usually visibility resulting in thin wings with high sweep back. Blowing of • Higher aspect ratio → lower CD,i → increased sustained high-lift devices would increase maximum lift but it also turn rate and better cruise/loiter performance. comes at the cost of unusable thrust. Note that high-lift Disadvantages of forward sweep: devices can also be used to improve maneuverability. • If the root stalls, vertical tail could be in the wake • Risk of divergence and new forms of flutter • Higher aspect ratio → higher CLα ⇒ III.MANEUVERABILITY – Higher gust sensitivity Requirements: – No aeroelastic relief • Superior transonic maneuvering is an important specifi- • Higher aspect ratio → lower CD,i → additional wave cation drag in supersonic conditions due to volume – Requirements on instantaneous maneuvering (pitch, roll, yaw rates) – Requirements on sustained maneuvers (turn rate, High-lift devices climb rate) Three categories of leading-edge devices: • For sustained maneuvers high specific excess power is • Alter leading-edge pressure distribution required • Alter the boundary layer (blowing and suction) – High lift, low drag, high speed, high thrust • Combination of both – Flight at high AoA leads to separation Leading-edge devices increase lift through increase of – Increase in drag, buffet and stability and control camber. They are most effective on sharp-nosed sections problems that are prone to separation. It is difficult to apply while • Result: degradation of combat capability: maintaining a smooth knuckle. Typical deflection is about 25 – Reduce pilot control and aiming accuracy degrees. Leading-edge devices cause a thicker wake over the – Full maneuvering capability is reduced trailing edge flap, this reduces the effectiveness of the trailing – Chance of stalling and spinning the aircraft edge flap. – Increase in drag reduces combat effectiveness

Kruger flap and slat without slot: Specific excess power • Increase wing chord or increase nose radius or both • Simple rotation about a hinge (Krueger flap) Specific excess power is a measure of the ability to (re)gain • Extension mechanism energy by accelerating or climbing. In level flight: Slat with slot: The normal load factor n can be computed by eq. 6. • Slat effect T = D0 + Di (1) – Reduces suction peak on main component 1 C = C + k(C )2 with k = (2) – Reduces adverse pressure gradient on main compo- D D,0 L πAe nent nW 2 1 T − D = kC2 qS = k qS with q = qV (3)2 • Circulation effect 0 L qS 2 – Slat in upwash of main wing T − D W ⇒ 0 = kn2 (4) – For Kutta condition at training edge of slat: circula- W qS tion (=lift) T − D 1 n2 = 0 q (5) • Dumping effect W kW/S ADVANCED AIRCRAFT DESIGN II SUMMARY 6

1 dE T − D ⇒ = V = P (19) W dt W S dh  V dV  ⇒ P = 1 + (20) S dt g dh √ r ρ √ dV V d σ V = V = V σ ; = − C (21) C ρ dh σ dh 0 √ dh  1 V 2 d σ  P = 1 − √C (22) S dt g σ σ dh dh PS ⇒ R/C = = √ (23) V 2 dt 1 √C d σ 1 − g σ σ dh Using specific excess power one can also determine the optimum energy climbs or make a plot of the airspeed vs. ¨ turn rate (so called doghouse plot¨. Flap scheduling Program flaps to automatically suit flight mode.

Buffeting Fig. 6. Derivation of specific excess power 1) Early formation of weak tip shock s 2) Overtaken by aft-moving shock from distorted pressure T − D 1 ⇒ n = 0 q (6) field at junction W kW/S 3) At higher Mach forward shock appears parallel and close to leading edge The relation between turn rate(ω), normal load factor(n) and 4) Forward shock moves inboard and intersects rear shock turn radius (R) is derived in eq. 12 outboard of intersection is a strong shock with a large pressure rise. This invariably causes flow separation V = ωR (7) W r 1 V ω = nW 1 − (8) g n2 W V ω = L sin φ ; φ = bankangle (9) g r 1 W = nW cos φ → sin φ = 1 − (10) n2 s g T − D 1 ω = 0 q − 1 in rad/sec (11) V W kW/S V V 2 1 R = = √ (12) ω g n2 − 1 In eq. 23 the relation between specific excess thrust and rate-of-climb (R/C) will be derived, starting with Newton’s law along the flight path.

W dV T − D − W sin γ = (13) g dt T − D V dV dh V dV Fig. 7. Flow over swept wings V = V sin γ + = + (14) W g dt dt g dt W 1 E = W h + V 2 ; E = T otal Energy (15) g 2 Vortex Lift E V 2 See figures 8, 9 and 10, that pretty much explains it. E = = h + (16) S W 2g dE 1 dE E dW S = − = P (17) Weapons vs. Maneuverability dt W dt W 2 dt S dW • Weapon capability determines aircraft agility require- = 0; weight constant (18) dt ments (both for attack and defense) ADVANCED AIRCRAFT DESIGN II SUMMARY 7

IV. AIR INTAKES Intake design criteria • Spillage drag • Internal performance (total pressure recovery) • Inlet/engine airflow matching • Flow distortion at compressor face – steady state – time-variant • Bypass flow Fig. 8. Vortex lift • Inlet bleed requirements • Interference with external flow • Stealth (radar detectability) • Boundary-layer diverter and bleed drag • Intake buzz and bypass drag • Flight and operational safety • Foreign object damage “The engine face average total recovery is of prime interest due to its direct effect on engine thrust.”

Fig. 9. Pressure distribution vortex lift Steady state distortion: • = pressure recovery pattern across the engine face • = felt by compressor blades as variation in velocity • Gun/cannon armament (ballistic unguided) is ”classical” solution • Results in vibrations of the blades • May cause stalling of blades on several stages • Early warning radar and guided missile development has eliminated high altitude penetration (SAM) • Can result in engine surge • Air-to-air missiles have longer range and maneuvering capability, plus higher speed than opposing aircraft (but disengaging is difficult or impossible) • Missiles have a minimum engagement distance, to lock on and stabilize flight plus a limited ”aiming cone” Fig. 11. Distortion • Cannon armament to supplement missiles • Long range engagement (BVR) identification not certain Dynamic distortion • Engagement may develop into close combat • = how the distortion/turbulence pattern varies with time • Missile capabilities improvements benefit from (short • High distortion levels result in low pressure recovery time) aircraft pointing capability • Multi-shaft bypass engine more susceptible to distortion • This may be traded against energy conserva- than pure jets tion/management Spillage drag • supermanoeuvrability • At high forward speed a low setting: stream tube smaller than inlet • Momentum loss of air that spills around the inlet = spillage drag. • Intake may be matched to flow conditions by variable geometry, blow-in doors etc. • Energy loss in bypass air, boundary layer bleed is pro- portional to mass flow velocity reduction Boundary layer bleed • Boundary layers impair pressure recovery • Goal: to remove fuselage and intake boundary layer • Means: use of boundary layer diverter • Result for fuselage boundary layer diverter → diverter drag (momentum lost by diverted flow • Result for intake boundary layer diverter: Boundary layer bleed drag (momentum lost from time they enter the intake until they leave the aircraft + exit door pressure Fig. 10. Effect of strakes on vortex lift drag) ADVANCED AIRCRAFT DESIGN II SUMMARY 8

• Goal: Make sure excess thrust due to higher pressure – Side by side vs. separated engine/inlets: transonic vs. recovery is not lower than additional drag supersonic performance • Boundary layer bleed is required for stable and undis- – Asymmetric engines have low spillage drag turbed engine intake flow – 2-D inlets have good pressure recovery with accept- Radar detectability able inlet drag – Nose intake (mainly used in early jet fighters, suf- • Diverterless inlets fered form high pressure losses due to wall friction, • Intake shaping very important for front radar cross- less flow distortion, no bl diverters necessary, no section large radar dishes at the time) Intake design features – Wing-root leading-edge intakes (small depth of in- • Intake size take face, rapid changes in cross section and low – Usually sized for high-subsonic speeds wetted area) – Excess airflow is diverted back to the freestream – Side intakes (induced by shape of nose, underbody, – Intake sizing should account for increased mass flow canopy, nose droop and fuselage camber, subject to due to engine development magnified AoA effects, need adequate handling of • Cowl lip shape fuselage boundary layer) – Shielded intakes (reduce intake AoA during ma- – Fixed profile neuvres, wing shielding improves pressure recovery, – Variable radius inlet massive flow separation in sideplate at high angle of – Suck-in doors (Alternative to high mass flow re- sideslip possible) quirement, they suppress separation without adding – Ventral inlet (fuselage is an efficient flow straightener thickness to the lips) when wider than inlet, low distortion, large pressure – Blunt lip avoids separation at low speeds recovery, magnified side-slip effect on intake inflow, – Blunt lip will cause shock wave and boundary layer nose wheel should be more aft, larger VT required) separation at high subsonic speeds due to spillage – Dorsal inlet (low RCS, bad high AoA performance) – Sharp lip causes flow separation at high angle-of- – Under wing inlet (under wing Mach number is attack and low M → can lead to distortion lower (precompression), high AoA capability, forces – Alternative: variable radius inlet/auxiliary intakes caused by flow spillage might actually improve the • Intake shape lift of the airplane, increasing L/D and easy access) • Sideplates Intake types • Intake boundary-layer management – Fuselage boundary layer (separated by splitter plate) • First generation of supersonic intakes: – Internal boundary layer (can be thinned by porous – sharp-lipped pitot intake surfaces or diverted by throat slot bypass – long subsonic duct (high internal friction) • Engine bypass system – large total pressure loss due to normal shock wave – Air that is captured but not accepted by the engine • Second generation: addition of conical spike (Mig 21) can be bled off using the incorporated boundary layer – Houses radar dish bleeding system – Improves supersonic pressure recovery (oblique – Much larger quantity of air than just the boundary shock) layer • Horizontal ramp inlet – If bypass/bleed are not used correctly a severe drag penalty can occur – Fuselage boundary layer diverter required – Long ramp lengths due to inlet aspect ratio (thicker • Intake duct length and shape boundary layer) – Air is decelerated by a (series of) shock wave(s) – Variable geometry capability in the ramp angle – Further diffusion required to decelerate to M =0.6 changes for mass flow regulation – Compatibility with area ruling (outside) and diffuser – Large side areas require sideplates to prevent side shaping (internal) spillage. Reduces stable mass flow range. BL growth – Result often S-shaped duct on sideplate. Flow separation off leading edges of – Duct length is trade-off between weight, distortion sideplate during sideslip conditions levels, diffuser losses (bl friction), (directional sta- – Ramp and throat boundary layer removal to mini- bility, example F-16) mize terminal shock/boundary layer interaction – to • Intake location improve subsonic diffuser performance and reduce – Engine intake to be optimized with : avoid distortion and turbulence. Aspect ratio chosen for disturbed flow, make use of precompression/flow best integration to aircraft configuration straightening – Small cowl lip area available for cowl suction (re- – No single configuration provided the best perfor- duction of spillage drag) but cowl drag is reduced. mance at all conditions – Cowl lip shaping for subsonic high angles-of-attack ADVANCED AIRCRAFT DESIGN II SUMMARY 9

– Possible inclusion of variable cowl devices to en- V. STEALTH hance inlet engine matching “The act of moving, proceeding, or acting in a covert way.” – Very good predictable angle-of-attack performance ∗ Subsonic: Ramp reduces inflow angles. Cowl lip Advantages of stealth: blunting can prevent internal flow separation. • Can penetrate highly hostile regions ∗ Supersonic: Shock system moves forward relative • Provides initial breakthrough by shock and surprise to cowl lips: maintains low distortion level of • Precision bombing intake air. Recovery may increase due to change • High Survivability in hostile conditions in effective capture area. • One mission, multiple targets • Half cone inlet • Cost effective in the long run – Variable geometry for mass flow regulation via trans- • High morale and confidence in the troops lating cone – Throat boundary layer removal necessary to mini- Linear changes in aircraft survivability produce exponential mize terminal shock/BL interaction to improve sub- changes in force effectiveness and aircraft attrition rates. sonic diffuser performance and reduce distortion both steady state and time varying. Susceptibility reduction: – Lower cowl angles required due to nature of conical • Threat warning flow • Noise jammers and deceivers – Long length of cowl lip perimeter available for lip • Signature reduction suction. Benefits to reduce spillage drag • Expendables – Structurally more efficient • Threat suppression – Splitter plate not required with proper diverter design • Tactics (which would minimize leakage off cone edge) – AoA performance: Vulnerability reduction: ∗ Subsonic: Lower cowl lip blunting required • Component Redundancy ∗ Supersonic: Asymmetric compression, increased • Component Location distortion. Shock pattern intersects plane of inlet. • Passive Damage Suppression Large degradation in recovery • Active Damage Suppression – Stable mass flow ratio change at Mach 2.0 is approx • Component Shielding 30 to 50% with an inlet design Mach number of 2.2 • Component Elimination/Replacement • Vertical ramp inlet Classification of aircraft signatures: – Variable geometry capability in ramp angle changes • Active: for mass flow regulation – Ramp and throat boundary layer bleed to minimize – Radar terminal shock boundary layer interaction and im- ∗ Airframe prove subsonic duct performance and reduce turbu- ∗ Engine Intake lence and distortion ∗ Weapons – Minimal side spillage areas due to aspect ratio (cho- ∗ Navigational Radar sen for best integration of aircraft configuration) • Passive: – Side plates eliminated to improve angle-of-attack – Infrared performance. Increase stable spillage range ∗ Fuselage – Large cowl/lip area – available for cowl suction – ∗ Airframe reduction of spillage drag ∗ Exhaust plume – Good angle-of-attack performance: ∗ Tailpipe ∗ Subsonic: Blunting of lower lip is required but ∗ Sun glint internal flow separation is prevented – Acoustic ∗ Supersonic: Shock pattern not greatly influenced ∗ Engine Parts by angle-of-attack. Small degradation in recovery. ∗ Engine Exhaust – Stable mass flow ratio at Mach 2.0: approx 10 to ∗ Airframe 30% with and inlet design Mach number of 2.2 – Visual Variable-geometry intakes ∗ Airframe • Moving cowl ∗ Engine Exhaust Glow • Extra chin intakes ∗ Canopy Glint • Rotating intake cowl ∗ Aircraft Lighting – Misc. ADVANCED AIRCRAFT DESIGN II SUMMARY 10

Shape Radiation Direction RCS ∗ Communication Sphere of diam. A any πa2 ∗ Countermeasures 4πA2 Flat plate (area A) normal to surface λ2 λ2 tan (δ) Cone (semi cone angle δ) Parallel to axis 16π πb2 Band Designation Nominal Frequency Ellipsoid (major axis 2a, minor axis 2b) Parallel to 2a a2 VHF 30-300 MHz Paraboloid with apex radius of p Parallel to axis 4πp2 UHF 300-1000 MHz λ2 tan δ Circular ogive (nose semi angle δ) Parallel to axis 4π S 2-4 GHz 2πaL2 Circular cylinder (length L and radius a) Perpendicular to axis λ C 4-8 GHz 4 90◦ 12πL X 8-12 GHz Trihedral (3 plane intersecting at ) Any angle between two faces λ2 K 12-18 GHz U TABLE II TABLE I RADARCROSSSECTIONS FREQUENCYBANDS

NOTE: RCS varies with frequency. Long range with low  2 2 1/4 frequency, long wavelength radars (resolution not so good), PR · GR · λ · σ Rmax = 4 (24) short range with high frequency, short wavelength (high (4π) · N · (S/N)min resolution)

Fig. 14. Reducing the RCS. Overall RCS reduction: Fig. 12. Relative size contribution to RCS • Reflection: – Minimise overall size of the aircraft – Clean external geometry having no protuberances or gaps – Internal weapons carriage – Highly swept leading edges – Eliminate cockpit transparencies – Use of composites – Use of passive onboard detection system (FLIR, IRST) – Use of radar screen on engine air intakes – Appropriate shaping of the intake lips and inlet ducts – Stealth aircraft must be low probability of intercept Fig. 13. Radar sources (LPI) • Absorption: Types of reflection: – Attenuating RAM • Diffuse reflection (rough surface) – Resonant RAM • Specular reflection (smooth surface) • Active Interference • Retro reflection (retroflecting foil of cat’s eyes reflector)

Radar reflection type: Planck’s radiation law: eq. 25 • Rayleigh Region (λ > α) 2π · h · c2  1  • Resonant Region (λ ≈ α) E(λ, T ) = 2 5 h·c (25) • Optical Region (λ < α) n · λ e n·λ·k·T − 1

4 Radar detection range, see eq. 24. Qemitted = σAT (26) ADVANCED AIRCRAFT DESIGN II SUMMARY 11

Plank constant h = 6.626 cot 10−34Js Boltzmann constant k = 1.3806 · 10−23J/K • Vertical ellipse = unstable in yaw Speed of light c = 229 792 458m/s • Horizontal ellipse = stable in yaw Refractive index n = 1 (for vacuum) • Flattened fuselage is longitudinally less stable (nose-up TABLE III pitch) CONSTANTS • Nose strakes allow for symmetric vortex formation • Apex of vortices is fixed • Strakes improve lateral/directional stability • Strakes prevent spinning Sources of IR signature: • Strakes deteriorate RCS • Emitting surfaces (engine casing, nozzle, exhaust plume, • the effect of strakes is also dependent on other other associated hot parts and airframe) components and can also be destabilizing in yaw. • Reflecting surface (sun glint off the airframe opaque surfaces and transparencies like cockpit canopy Nose shape affected by radar • Primary geometric factors affecting radar performance: VI.FUSELAGE DESIGN – Location of a pitot-static boom, nose strakes (and Functions of the fuselage is to accommodate: sometimes AOA and angle-of-sideslip vanes) adja- • crew cent to the radome. • communications and navigation equipment – High fineness ratio, resulting from aerodynamic • the flight control system shaping of the nose for low drag. • search and fire-control systems – Aerodynamic shaping of the nose cross-section for • large proportion of its fuel load good high-AOA directional stability • Engine(s) • Trade-off between • Components of the – Low drag • Gun + ammunition – Excellent high-AoA characteristics • Missiles, bombs, flares – Acceptable radar performance

Nose and forward fuselage Forebody affected by crew Forebody shape is driven by • Lage fineness ratios desired for supersonic flight, this • Cockpit visibility requirements which usually govern requires low cross-sectional area → no more side-by-side forebody camber. cockpits, but rather tandem. • High-AoA handling which influences the length, cross- • High visibility for the pilot improves survivability. sectional shape and application of nose strakes • Requirement for radar and laser-ranging installations in- fluece the nose size and shape Center Fuselage • Crew accommodation, including cockpit canopy design, Center fuselage accommodates: governs the cross-sectional area • Main ducts for engines • Fuel tanks Forward camber: positive camber generates a negative • The main undercarriage (optional) pitching moment, so reducing the forebody camber will • Armament bay (optional) reduce the horizontal tail download required. • Ejection units for stores (optional) • Pipes, cables and wiring Forebody vortex flow: • has dominant effect on stability in post stall (high AoA Some requirements in the design of the center fuselage are conditions) space, accessability (especially for engines) and vulnerability • vortices are shed from the nose of the airplane of systems. • fin subjected to wing wake and vortices shed by the forebody • Internal engines: when engine grows, more space is • it is influenced by nose fineness ratio (large yawing required. moments if large fineness ratio), bluntness, cross- • Podded engines: sectional shape and the use of nose strakes (aka spin – Engine growth easily realized strakes/strips) – No need for fuselage boundary layer diverter – More wetted area Forebody effect on stability – Heavier structure (of pods) • A well-designed forebody can also contribute to positive – Less wing weight (inertia relief) directional stability at high angle of attack – Less intake weight (no S-duct) • Requirement = stable separation – OEI condition is more critical ADVANCED AIRCRAFT DESIGN II SUMMARY 12

– Better longitudinal stability (pancake area aft) – destabilizing fuselage – Better directional stability (mode side area) – requirement to suppress sideslip rapidly – lateral stability requirements Area rule: Any two bodies which have the same area • Result distribution, will experience the same amount of wave drag – Difficult to predict required fin size for directional independent of their actual shape. stability External store asymmetry: • “Sears” bodies describe the optimum shapes for minimum • Imbalances occur when stores are released asymmetri- wave drag. cally • “Sears-Haack” body = minimal drag for given length and • Mass asymmetry causes rolling moment volume. – Can be balanced by aileron deflection • “Von Karman ogive = minimal drag for given length and – Aileron deflection produces variable rolling moment cross sectional area. during high-g maneuvers – Required lateral moment can be relieved id airplane Differential area ruling is allowed to sideslip but this requires sufficient • Favorable lift interference is created by differential area directional stability ruling • Missile firing generates asymmetric flow field, this can • Requires lower AoA to attain certain CL generate a yawing moment • Reduces drag due to lift (remember this is proportional to C tan α) L Compressibility effects: • 5% increase in sustained turn rate at M=1.2 • Fin usually becomes less effective at higher Mach • Favorable pitching moments with reduced trim drag numbers Area ruling conclusion: • Lowest body drag: At high angles of attack the low energy wake (wake of – Large fineness ratio the wing/forebody) can immerse the fin, this can result in – Proper area ruling yaw-off due to directional instability. • Issues: At subsonic speeds it is mainly the fin height that influences – Large pitching moment of inertia stability (not so much the area). However, because of fuselage – Large yawing moment of inertia vortex interaction with the vertical fin, at high angels of attack – Small rolling moment of inertia and at yaw angle, a higher fin can cause reduced stability. – Hazard of inertia cross coupling Twin fins Rear Fuselage • Can reduce required tail height (→ less aeroelastic penal- Requirements: ties) • Mutual interference at subsonic conditions reduces effec- • Have to be compatible with large tail plane angles tiveness and could render a single fin more effective • Minimize leakage between HT and fuselage through gaps • Most beneficial in supersonic conditions; beyond a • House HT actuators (pivot axis of HT should be close to certain Mach number there is no mutual interference, its AC for actuator sizing) the Mach lines do not interfere with the other • House engine (usually in the back to reduce structural heating and acoustic effects) Reasons for applying twin fins on the F-14:

VII.FINSAND • “End plating” effect if the twin verticals results in far more effective horizontal tail control. Functions: • Twin verticals provide a more constant value of Cnβ and • Balance in asymmetric flight high Cnβ for improved Dutch roll characteristics. • Ensure maneuverablilty • control redundancy for combat survivability. • Provide directional stability (weathercock function) • Better infrared stealth due to exhaust shielding. • Spin prevention/recovery • No spine boom required to mount fin between engines • Reduced height makes them less sensitive to flutter ∂Cn ∂β must be positive. • Larger distance form centerline required less heavy struc- ture → lower temperatures, less noise. Fin area requirements: • Lower height means less hanger space required • Directional moments generated by – Destabilizing forebody Reasons for not choosing twin fins on YF-16: – Stabilizing fin • Flow separations from both forebody and wing at high • Fin size dominated by AoA interacted adversely with twin verticals. ADVANCED AIRCRAFT DESIGN II SUMMARY 13

• Under certain combinations of α and β visible buffeting • Aeroelastic distortion reduces effectiveness of rudder at of the tails occurred. high dynamic pressure • Single fin reduced friction drag due to lower required • ⇒ high aspect ratio rudders not used; wide, swept wetted area. rudders are used as well as all-moving fins

The fin is usually placed as far aft as possible (taking into account area ruling and interference between horizontal and VIII.NOZZLESANDAFTBODIES vertical tail at high AoA). Nozzles are needed to control the expansion of exhaust Canting the fins: gasses, by preventing uncontrolled expansion one can achieve • Reduce rolling moment. increased gross and net thrust. • Reduce radar cross-section. • Reduce adverse interference with forebody vortices, wing vortices and vortices. • Toe angle may be required to reduce vortex interference.

Fin shape: • High aspect ratio is beneficial for effectiveness • Low aspect ratio is beneficial for stall behaviour Fig. 15. Area ratio changes by a factor of 1.7 to 1.8, theoretically something • Alternative: dorsal fin like 3 would be best but that would result in very large areas and lots of drag. • Raked fins on Russian fighters: increase in flutter speed Ventral fins: Off-design operation, low supersonic speeds: • Oppose roll due to sideslip • Nozzle expands flow below back pressure. • Is destabilizing in roll: • Corrective shock wave occurs. – offsets the dihedral effect • Shock-induced separation follows. – good for Dutch roll characteristics • Effective loss not as high as in ideal case (normal shock), but the total loss is still substantial • Low aspect ratio • High structural stiffness • At low AoA might interfere with fuselage stores Off-design operation, low subsonic speeds: • Divergent part acts as subsonic diffuser Rudders • Velocity decreases downstream • pressure increases downstream Design factors • Adverse pressure gradient can cause unstable separation • Crosswind landing (often most critical) • Causes the jet stream to attach to one side and then the • High AoA flight (spin recovery) other • Asymmetric stores • Causes violent vibrations • Asymmetric thrust (engine failure, engine unstart) • Transonic effectiveness Thrust vectoring is the manipulation of jet exhaust such that the resultant reaction forces augment or replace those forces High AoA flight normally generated by the aerodynamic control surfaces of • The rudder effectiveness is determined by the wing the aircraft. planform and its stall pattern (position of the stalled wake Thrust vectoring can give you increased range (no need for w.r.t. the rudder). Also (part of) the rudder can be blanked trim-deflection of control surfaces), improved agility and by the wake from the horizontal tail. If the rudder is not better survivablilty. used for spin recovery however that is no problem, • Right yaw + right aileron: Axis-symmetric nozzles

– combined with low Cnβ at high AoA: sideslip to the • High strength/weight left is induced • Easier to cool – combined with low Clβ: a roll to the left in incurred • Less leakage between upper/lower ramps and sidewalls (roll reversal) • If aileron reversal occurs switch to rudder-only roll 2D thrust vector control control beyond a certain AoA, this is automatically done • Reduced complexity via Aileron-Rudder Interconnect (ARI) • Easy integration with aft fuselage • More effective TVC Transonic effectiveness • Lower IR signature • Compressibility effects reduce effectiveness of hinged control surface in transonic and supersonic speeds Thrust vectoring conclusion ADVANCED AIRCRAFT DESIGN II SUMMARY 14

• Only increases maneuverability at low speeds • Can replace (some part of) control surfaces and could reduce weight; no RCS benefit, small benefit in endurance • Can allow for lower landing speeds if TVC can roll the airplane (limited by impaired pilot visibility)

Afterbody contours: How to reduce subsonic boattail drag? • Expansion of the nozzle until its area coincides with that of the nacelle. This can eliminate base drag but it causes jet over-expansion and internal losses. • Careful design of the boat-tail. A minimum boat tail angle at 15◦ has been suggested, the maximum angle is dominated by boundary layer fatigue. If the angle is smaller than 10◦ this will result in extra length, with extra skin friction and excess rear fuselage weight. (possible solution is the addition of a nacelle-to-nozzle fairing (F- 16)) • Operation of the nozzle at limited area ratio

Boundary layer separation on the boat tail might cause: • Excessive boat tail drag • Violent buffeting Reduction in boat tail angle: • Causes flat base area • Clean boundary layer detachment at the rim of the fuselage • Does not result in too high drag when close to an exhaust

Interference drag • Change in boat-tail pressure w.r.t. tail without nozzle and jet + nozzle drag + change in gross thrust due to the external flow field • interference drag reduces with increased nozzle spacing (optimum spacing to minimize total drag s/d = 2.5)