AIRCRAFT STRUCTURES: RIGID AND ELASTIC

1. AIRCRAFT STRUCTURES:RIGID AND ELASTIC Dr. John Valasek Aerospace Engineering Texas A&M University AERO 401 November 1999

2. INTRODUCTION early motivations • The main factor that governs the choice of materials and structural form is the ratio of the load on the structure to its dimensions. • mission type and speed • Very early aircraft operated at low speeds, and therefore loads were low in relation to aircraft size. Wing loadings were typically 5 - 10 psf. • best option was to concentrate compression loads into a few small rod-like members and diffuse tensions into fabric and wires • Low power engines of the time made structural lightness an expedient • wood and fabric were best choice, and simple to obtain • aircraft of similar dimension were less than the weight of comparable modern ones • metals were entirely out of the question • Biplanes were prevalent because early monoplanes suffered from catastrophic structural failures (probably caused by aeroelastic effects which were unknown at the time). • WWI dogfight load factors could be as high as 4g

3. INTRODUCTION modern motivations • For high speed flight, the main factor that governs the choice of materials and structural form is the high temperature environment caused by kinetic heating in sustained supersonic flight. • Except for one or two exceptions, the top speed of fighter aircraft have traditionally been limited not by aerodynamics or propulsion but by the choice of materials. • without advances in structural efficiency the performance improvements due to advances in aerodynamics and propulsion would not have been realized • Existing fighter aircraft as a rule do not have long supersonic endurance, and so have metalic leading edges (for reasons of rain and birdstrikes). • The proposed U.S. High Speed Civil Transport (HSCT) is critically dependent on advanced structures and materials technology.

4. Thick Skinned Jets Deltas Wood Biplanes Metal Monoplanes WING LOADING fighter aircraft trends (1910 - 2000) • Wing loadings based on maximum takeoff weight. • The great rise in wing loading occurred during the 1930’s and 40’s. • The generation of fighters with thick skins lessened the trend slightly. • Note the difference in F-16 takeoff wing loadings: • F-16A air superiority • F-16C multi-role

5. Moraine-Saulnier Type N “Bullet” WING CONSTRUCTION the early years (1900 - 1918) • Fabric covering wooden spars. • Load carried by internal structure plus bracing wires. • Typical of WWI aircraft. • Load bearing members are positioned near aerodynamic surfaces where the stresses are highest. • Upper surface in compression, lower surface in tension. Stresses near the neutral axis are low and lightening holes can be used. • Susceptibility to structural failure due to wood rot. • Buckling of wings in flight called a “striptease” in the vernacular of the period.

6. Military Wing Sopwith D.1 No. 243 Squadron STATIC LOADS TESTING 1920 Determining ultimate flight loads by testing to destruction

7. Hawker Hurricane Mk. XII WING CONSTRUCTION the inter-war years (1919 - 1938) • Built up steel spars with wood reinforcement, covered with fabric. • Warren type truss. • Load carried by internal structure plus bracing wires. • Intended to be the “best of both worlds” in terms of greater structural strength due to inclusion of steel, and lower cost, ease of manufacture, and ease of maintenance due to fabric covering. • Ended up being “worst of both worlds” • mix of steel and wood not as strong as steel alone • fabric unable to withstand higher speeds permitted by stronger structure

8. Messerschmitt Me 262 Sturmvogel WING CONSTRUCTION WWII to Korea and after (1939 - 1955) • A major conceptual breakthrough: most of the structural load is carried by the externalstructure. • Semi-Monocoque construction • the thin skin can easily handle tension • to handle compression without buckling, the skin is attached to the spars and stringers • Stressing the skin results in an even higher load carrying capability. • total result is a structure very stiff in bending. • requires mechanical fasteners (rivets). • permits higher speeds / lower drag. • Discovered in 1925 by Dr. H. Wagner, termed the ‘Wagner Theory of the Diagonal-Tension Field Beam,’ • Standard construction type today.

9. WING WEIGHT fighter aircraft trends (1930 - 1980) • Normalized to P-51 baseline span (accounting for planform, section, materials). • Modern jet wings are much lighter than 1940’s prop wings • P-51 14.5% WTO • F-15A 3% WTO • If modern wings had to be built using 1940’s technology, they would virtually be solid aluminum alloys or steel. • Structural efficiency has improved greatly with time.

10. McDonnell F-101A Voodoo WING CONSTRUCTION supersonic to post Vietnam (1955 - 1975) • High transonic and supersonic flight speeds mandated wings with • very low thickness ratios • large bending strength • sweepback torsion • thicker skins and therefore more structural material. • Solid wings were one answer (F-104). • A better solution is integral wings • skin and stringers are machined from a single large piece of material • eliminates mechanical fasteners • good surface finish (low drag) • “Wet Wing”; no bladders, but integral: • fuel tank • torque box • skin • significant increase in fuel volume • structural synergism

11. Sukhoi Su-27 Flanker INTERNAL FUEL LOAD fighter aircraft trends (1945 - 2000) • Comparison of integral tanks and bladder tanks. • For the same area, integral tanks offer greater capacity. • Notable aircraft: • F-101A fuselage fuel • F-15E conformal tanks • Su-27 overload condition Integral Tanks Deltas Bladder Tanks Pre - 1955

12. WING CONSTRUCTION contemporary (1976 - ) • The quest to save weight while still retaining good mechanical properties. • Concept: reduce structural mass by reducing material density, instead of increasing mechanical properties like • strength • stiffness • toughness • For most materials: • 10% strength increase, 3% weight reduction • 10% density reduction, 10% weight reduction • Execution is usually in the form of various types of alloys and composites. • Drawbacks include • cost • difficulty in manufacturing • undesirable aeroelastic effects such as reduced roll rates and aileron reversal

13. Saab JAS 39 Gripen STATIC LOADS TESTING 1998 Non-destructive testing including accurate measurement of deflections

14. TOTAL GROSS WEIGHT REDUCTION projected Source: Aerospace America, November 1997

15. THE COMET STORY (1) 1949 A New Era Begins The DeHavilland D.H. - 106 ushers in the jet age in commercial air passenger transport DeHavilland D.H.-106 Comet 1949 48 pax 490 MPH 3540 nm

16. THE COMET STORY (2) 1953 - 1954 Tragedies • Five aircraft are lost • two due to stall at takeoff • three inflight, due to “unknown” causes • BOAC Comet Yoke-Peter, serial G-ALYP, (the first Comet I in scheduled service) crashes off the island of Elba in the Mediterranean Sea, 10 January 1954. 35 pax plus crew are lost. • South African Airways Comet crashes off the island of Stromboli in the Mediterranean Sea, 8 April 1954. 14 pax plus crew are lost. • Deep sea salvage using sonar and underwater television cameras is used for the first time to locate aircraft wreckage.

17. THE COMET STORY (3) 1955 The Cause Revealed • The Particulars • pressurized cabin • multiple pressurizations / depressurizations • square windows • The Mechanism • crack propagation • The Result • structural failure resulting from repeated loading/unloading cycles • The Phenomena • Cyclic Fatigue

18. Boeing 367-80 1954 118 pax* 582 MPH 3530 nm Douglas DC-8 1958 132 pax 600 MPH 3550 nm THE COMET STORY (4) America responds

19. THE COMET STORY (5) the lead is lost for good • The improved “safe version” Comet 3 (1955) and improved range (transatlantic) Comet 4 (1958) are offered. • In 1958 the Comet 4 begins the very first regularly scheduled transatlantic jet service. • westbound flights still had to refuel at Gander, Newfoundland • One year later, the DC-8 and B707 firmly captured the market due to higher speed and significantly larger passenger capacity. • Comet 4: 76 pax at 500 MPH • B707: 176 pax at 600 MPH • Comets are eventually sold to the Royal Navy as Nimrod AEW aircraft.

20. Boeing V-22 Osprey FATIGUE TESTING ensuring long term structural integrity • V-22 design life is 10,000 hours, or 20 years of flying ops. • Airplane andhelicopter induced loads will be encountered. • takeoffs • landings • airplane and helicopter maneuvers • rough field and shipboard operations • ground maneuvers (braking and taxiing) • For acceptance, structural integrity of airframe is tested to multiple lifetimes. • Two for low-cycle loadings (20,000 hrs), three for high-cycle loadings (30,000 hrs) • Minimum 7,000 hours in airplane mode, 3,000 hours in VTOL mode. • No damage at 4g, 310 kts, and 2.8g, 345 kts. • At end of first test lifetime, airframe is disassembled and inspected. Source: Aviation Week & Space Technology, 20 April 1998

21. THE ELASTIC AIRPLANE fact or fiction?

22. AEROELASTICITY when flexible structure meets dynamic pressure Source: Air International, Vol. 52 No. 3, March 1997

23. Rutan Voyager ELASTIC AIRCRAFT practical considerations • All aircraft are elastic to some extent. • The designed-in level of airframe elasticity is dictated by: • operational requirements and constraints • aerodynamics • materials • economics • safety, e.g “bend but don’t break” • Some aircraft types are significantly more elastic than others: • Aircraft which are generally rigid • fighters F-15 Eagle • general aviation Cessna 172 • homebuilts made of conventional materials Thorpe T-18 • Aircraft which are generally elastic • supersonic cruise Concorde • large and long range transports and bombers Boeing 777 • homebuilts made of composite materials GlassAir

24. Boeing Model 707-320B Boeing Model 707-320B AEROELASTIC EFFECTS (1) steady-state stability derivatives • Example: Boeing Model 707-320B • Weathercock Stability • Elastic stability derivatives are a strong function of dynamic pressure and therefore speed and altitude. • Compared to the rigid aircraft: • elastic weathercock stability has essentially equal yet opposite slope for 0.1  M  0.9 • elastic weathercock stability is reduced 85% at M = 0.9 Source: Airplane Flight Dynamics and Automatic Flight Controls, Part II by J. Roskam

25. Model the horizontal tail as a flexible cantilever beam: Under a vertical load Lh the fuselage will produce an elastically induced angular deflection KLh. An up load produces a negative change in horizontal tail angle-of-attack. The total aerodynamic load is: Boeing Model 707-320B Note that Lh is a function of itself. Solving for this load: At high dynamic pressure the loads decreases because the denominator grows large. Converting to a pitching moment coefficient and differentiating with respect to e, AEROELASTIC EFFECTS (2) aft fuselage bending • Example: elevator effectiveness degradation due to fuselage flexure. Source: Airplane Flight Dynamics and Automatic Flight Controls, Part II by J. Roskam

26. MODELING AEROELASTICITY perturbed-state stability derivatives • Analytical derivatives are obtained by influence coefficient methods. • Aerodynamic [A] • rigid body Each element aij is the aerodynamic force induced on panel i as a result of a unit change in angle-of-attack on panel j. The column of aerodynamic forces is related to [Aij] and the airplane angle-of-attack distribution Source: Airplane Flight Dynamics and Automatic Flight Controls, Part II by J. Roskam by . Converting to pitching moment coefficient and taking the derivative with respect to pitch rate, gives the rigid body pitch damping derivative.

27. Undeformed or Jig Shape Elastic Equilibrium State JIG SHAPE (1) equilibrium states of elastic aircraft • It is assumed that: • The aircraft is held in its elastic equilibrium shape by an elastic equilibrium load distribution (gravity, aerodynamic, thrust). • The aircraft is elastically deformed in the equilibrium state. • strain energy is “pent up” in the structure • While under equilibrium loads, the center of gravity does not correspond to a specific point on the structure of the airplane. • When equilibrium loads are removed, the C.G. is a fixed point on the structure of the aircraft in its undeformed or jig shape. Source: Airplane Flight Dynamics and Automatic Flight Controls, Part II by J. Roskam

28. JIG SHAPE (2) equilibrium states of elastic aircraft • Elements of a calculated “jig shape matrix” must be translated into “jigging points” for the assembly jigs. • Determination of the jig shape is usually performed by computer. • Computer controlled laser-guided alignment is used during assembly.

29. ELASTIC AIRCRAFT summary • Multiple and simultaneous aeroelastic behaviours are typically encountered: • aileron reversal • wing divergence • loss of longitudinal control power due to aft fuselage bending • Aeroelastic effects on stability and control derivatives are usually significant and always vary strongly with flight condition. • Steady-state and perturbed state stability and control derivatives are fundamentally different for elastic aircraft: • inertial effects due to mass distribution invoke elastic deformations, altering the aerodynamic loading • Elastic aircraft must be designed, manufactured, and built to a jig shape to achieve a specific desired cruise shape under flight loads. • Many analytical modeling techniques exist of varying complexity and accuracy.