«by: Alexander M. Benoliel Thesis submitted to the faculty of the Virginia Polytechnic Institute & State University in partial fulfillment of the ...»
The airport noise problem also led researchers to study aerodynamically efficient, low speed, planforms to reduce required engine thrust, and thus takeoff noise. One of the reasons cited for canceling of the SST program in 1972 was community noise.3 It is difficult to generate low drag lift for takeoffs and landings using a slender wing. This presents a challenge in designing an efficient high lift system capable of meeting the requirements of the HSCT.4 The Concorde, with its slender wing, relied primarily on the lift generated by the leading edge vortex at high angle of attack to generate enough lift at low speeds. However, the drag penalty associated with the vortex lift resulted in a severe noise problem due to the required thrust for takeoff.4 This work describes an investigation conducted to review previous research done on high-sweep, low aspect ratio planforms and the current and past prediction methods developed. The goal of this work was to compile and assess past research, with emphasis Page 1 on understanding the pitch-up problem, determining the effectiveness of various leading and trailing edge flap configurations, understanding the benefits and shortfalls of tail, tailless, and canard configuration, developing an analysis method to estimate pitch-up using current linear aerodynamics codes, and applying this method to the determine the trim requirements of slender wings at high lift coefficients.
1.1 Past Research Research on High Speed Civil Transport (HSCT) aircraft has been conducted for the past thirty-five years in the SCAT, SST, SCAR (SCR), and HSR programs.3 Two configurations from the SCAT program that carried over to the SCAR program for further wind-tunnel testing were the SCAT-15F and SCAT 16. The SCAT-15F (Fig. 1a) was the fixed wing version of the SCAT-15 with leading edge sweep angles of 74˚, 70.5˚, and 60˚.
The SCAT-16 (Fig 1b.) was a variable sweep configuration similar in design to the Boeing 2707-100,3 which was Boeing’s initial entry in the SST competition. Problems with this wing-tail configuration were exhaust scrubbing and acoustic noise/fatigue on the passenger cabin and aft fuselage; and pitch-up in both swept and unswept wing positions. The winning 1967 Boeing SST proposal 2707-200, while still a variable sweep wing design, reverted to a high sweep, low aspect ratio planform when the wing, in the most aft swept position, was integrated and locked onto the horizontal stabilizer. The four turbojets were mounted beneath the horizontal stabilizer exhausting behind the trailing edge. The resulting high-speed configuration was a classic slender delta with a long, overhanging forebody.
Although Boeing won the SST contract with the 2707-200, they revised the design in 1969 into the fixed wing 2707-300 because there were overwhelming technical problems associated with the variable sweep wing design.5 These problems included aeroelastic effects due to the long fuselage, the need for a canard to meet takeoff rotation requirements, low values of lift-to-drag ratio for loiter due to outboard panel stall, and main landing gear placement in relation to engine location.
Figure 1. - SCAT Program developed configurations (not to scale).
Page 2 During the SCAR program wind-tunnel tests were conducted to evaluate vortex flaps, blown flaps, and the effects of tail and engine placement6-19. A collection of key work is included in Tables B1 and B2 in Appendix B. Many of the early models tested were variations of the SCAT-15F design, and the Advanced Supersonic Technology (AST) series configurations evolved from this work. Other configuration design studies available on HSCT type concepts developed in the past are documented in reference 5 and 20 through 28.
Page 3 2. Aerodynamic Pitch-Up
Pitch-up is defined herein as an abrupt change in slope of the CM(α) curve such that the slope of the CM(α) curve after the pitch break is greater than it was before the pitch break. The magnitude of the change in slope of the CM(α) curve defining the pitch-up varies depending on the configuration. For some configurations it is mild and may be difficult to identify. An example of pitch-up is shown for a 71˚/57˚ sweep wing, tested by Yip and Parlett19, in Fig. 2. The pitch break occurs at an angle of attack of about 6˚, corresponding to a lift coefficient of about 0.24. The figure includes a comparison with a vortex lattice numerical prediction method developed by Carlson, et al.29 Note that the nonlinear pitching moment occurs well within the operating regime of the aircraft and theory fails to predict it. Also, cranked arrow wings, such as the one shown in Fig. 2, are much more susceptible to pitch-up compared to pure delta wings.
Figure 2. - Lift and pitching moment for a McDonnell Douglas 71˚/57˚ sweep cambered and twisted cranked arrow wing (ref.
2.1 Theorized Reasons for Pitch-Up For typical HSCT-class wings pitch-up is a result of the forces generated by the leading edge vortex inboard, together with flow separation and vortex breakdown on the outer portion on the wing. The strong effects of the leading edge vortex, and the loss of lift on the outboard wing sections due to flow separation, causes the center of pressure to move forward producing the pitch-up behavior. This is similar to the flow phenomenon Page 4 encountered on high aspect ratio swept wings.30 The specific flow phenomenon which leads to the pitch-up distinguishes wing concepts. Some researchers believe that the pitchup is a result of the vortex breakdown at the trailing edge, which progressively moves forward with angle of attack. It is likely, however, that the vortex may move away from the surface and lose influence before vortex breakdown occurs. This was identified in Lamar's discussion of experimental results.31 It is more plausible that pitch-up is due to a combination of effects including vortex breakdown, but primarily due to outboard flow separation. It is important to identify if the pitch-up is dominated by the outboard flow separation or the strong inboard leading edge vortex, a function of the configuration.
Early work32 was done to predict which types of configurations were susceptible to pitch-up to provide guidance for use in preliminary design. This work produced the wellknown DATCOM design criteria for acceptable sweep and aspect ratio combinations. This method will predict if pitch-up will occur, although it does not define the angle of attack, or the lift level, where it will occur. It is also difficult to apply this method to cranked arrow planforms, in which more than one sweep angle is relevant.
A leading edge vortex on slender wings is created when the flow separates at the leading edge and then reattaches downstream on the surface, creating an area of low pressure above the leading edge on the upper surface (Fig. 3). As the angle of attack is increased, the core of the main vortex moves inboard11,34 and remains coherent up to larger angles of attack for higher sweep wings as shown by Wentz and Kohlman.35 Figure 3. - Leading edge vortex features on highly swept wings (ref. 33).
For a cranked arrow wing, two vortex systems may be formed due to the leading edge flow separation on each wing section. The inboard vortex can extend into the aft Page 5 portion of the outboard wing section. It also induces an upwash on the outboard wing section. The flow incidence angle on the outboard portion of the wing is considerably higher than the aircraft angle of attack due to this upwash. At low angles of attack, the vortex flow on the outboard wing section increases the longitudinal stability. This result is due to the fact that the outboard wing section is aft of the center of gravity, thus contributing a nose down moment. As the angle of attack increases, the outboard vortex system breaks down. At the same time, the inboard system moves further inboard, thus unloading the outboard wing section, as shown by Coe, et al.8,10 Rao36 also studied this outboard vortex breakdown in a test of a 70˚/50˚ sweep flat cranked delta wing. Through oil flow and smoke visualization, he showed the onset of vortex breakdown and flow separation on the outer wing panel at angles of attack as low as 5 degrees. This loss of lift on the outboard portion of the wing in conjunction with the strong inboard leading edge vortex causes pitch-up on slender arrow wings. The particular wing concept determines if flow separation on the outboard wing panel or the inboard leading edge vortex will initiate the pitch-up.
2.2 Influence of Geometry on Pitch-Up Several factors affect the pitch-up behavior of cranked arrow wing planforms. The introduction of a trailing edge notch places greater demands on the wing leading edge region. This effect is clearly seen in arrow wings with increasingly large trailing edge notches as shown in Fig. 4 taken from Poisson-Quinton.37 As the angle of attack is increased, the wing-tips become unloaded and the vortex core moves inboard. With the large trailing edge notches, the vortex has less area aft to affect, causing a destabilization in the longitudinal stability. Note that the pure delta wing does not encounter pitch-up.
The size of the trailing edge notch of an arrow wing can dramatically affect the pitchup behavior of highly swept wings. It was shown by Grafton38 that the addition of a trailing edge extension on a modified arrow wing planform (Fig. 5a) reduced pitch-up (done as part of the F-16XL planform development program). Although this modification did not change the angle of attack for pitch-up, it did reduce the severity of the pitch-up, as shown in Fig. 5b. Grafton also found major effects resulting from a leading edge notch on the same model. Here, the leading edge notch weakened the leading edge vortex38, resulting in a reduction of the severity of the pitch-up. This result demonstrates the possible sensitivity of pitch-up and lift characteristics to small planform changes brought about if these small changes produce a fundamental change in the leading edge vortex.
(a) Model diagram with trailing edge (b) Lift and pitching moment data for several extension, leading edge notch, and configurations (ref. 38) wing fence.
Figure 5. - Modified F-16XL predecessor model Much like the leading edge notch, the shape and incidence of the leading edge can affect the vortex lift.
Increasing the leading edge radius has the effect of improving the longitudinal characteristics by retarding the formation of the leading edge vortex. 12,13,16,33 This also reduces the vortex lift, as shown in Fig. 6. The local angle of attack of the leading edge seems to have the greatest effect on the aerodynamic characteristics with respect to the pitch-up. To minimize the formation of the vortex, it is desirable to deflect the leading edge Page 7 such that the leading edge incidence relative to the local flow angle of attack at each spanwise station is zero. Then, at angles of attack above this condition, the vortex will be formed uniformly. The leading edge can also be shaped and deflected such that the leading edge vortex is maintained on this surface.39 Applying this concept to a leading edge device results in the so-called vortex flap. Generally used for transonic maneuverability, deflecting the leading edge allows for the development of vortex lift while recovering some of the leading edge-suction to reduce drag.39 The vortex flap shape and wing camber must be optimized for minimum lift-induced-drag to be effective. The effects of leading and trailing edge flaps will be discussed in more detail in the following section.
Figure 6. - Effect of leading edge radius on lift and pitching moment on the SCAT-15F (ref.
16) Planform effects, such as the outboard wing sweep, were studied by Hom, Morris, and Hahne.40 Hom, et al, theorized that the pitch-up was a result of flow separation on the outboard wing panel, a function of the spanwise flow, and vortex breakdown. Four models were tested with a 70˚ sweep inboard wing section and varying outboard wing sweeps ranging from 60˚ to -20˚. They found that the lower sweep outboard wing panels encountered less spanwise flow and thus, the flow remained attached on the outboard panels. As the outboard sweep angle was increased, the flow on the outboard wing section separated and became dominated by a leading edge vortex on this section. The angle of attack for pitch-up was found to be unaffected by the outboard wing sweep. However, helium bubble flow visualization techniques showed that the cause of the pitch-up varied for the cambered and uncambered wings. It was found that the pitch-up for the uncambered wings was due to vortex breakdown at the trailing edge. When leading and trailing edge flaps were Page 8 deflected (to postpone the formation of the leading edge vortex) the pitch-up was a result of basic flow separation on the outboard wing section and not vortex breakdown.
Some of the configurations developed such as the SCAT-15F, the AST-100 and AST-200 series, and the General Dynamics F-16XL, incorporate the use of vertical fins or fences located outboard on the wings (generally placed at the crank location). Grafton38 found that the effect of fences located just inboard of the wing crank on a predecessor of the F-16XL (Fig. 5a) reduced the lift and created a slight improvement in the pitch-up at high angles of attack as shown in Fig. 5b. In a later test of a similar model, Grafton and Nguyen41 found that the slope of the pitching moment curve after the pitch-up increased with the addition of the wing fences at moderate angles of attack (decreasing longitudinal stability), and improved the longitudinal stability slightly at high angles of attack.
Lockwood13 tested a modified SCAT-15F model and found that vertical fins (whether located at the crank or inboard of the crank) reduced the longitudinal stability as well as the lift. This result is contradictory to Grafton and Nguyen’s findings. Both researchers agreed, though, that the effects were likely due to the loss of vortex influence on the outboard wing section and vortex breakdown due to the presence of the fences. Another effect of the fences or fins in these tests was to cause the inboard vortices to break down symmetrically in side-slip, thus improving lateral stability.13,38