Hypersonic control effectiveness studies on delta wings with trailing-edge flaps

Imperial Users onl

Lift augmentation for highly swept wing aircraft

A pair of spaced slots, disposed on each side of an aircraft centerline and spaced well inboard of the wing leading edges, are provided in the wing upper surfaces and directed tangentially spanwise toward thin sharp leading wing edges of a highly swept, delta wing aircraft. The slots are individually connected through separate plenum chambers to separate compressed air tanks and serve, collectively, as a system for providing aircraft lift augmentation. A compressed air supply is tapped from the aircraft turbojet power plant. Suitable valves, under the control of the aircraft pilot, serve to selective provide jet blowing from the individual slots to provide spanwise sheets of jet air closely adjacent to the upper surfaces and across the aircraft wing span to thereby create artificial vortices whose suction generate additional lift on the aircraft. When desired, or found necessary, unequal or one-side wing blowing is employed to generate rolling moments for augmented lateral control. Trailing flaps are provided that may be deflected differentially, individually, or in unison, as needed for assistance in take-off or landing of the aircraft

Low-speed wind tunnel study of longitudinal stability and usable-lift improvement of a cranked wing

An exploratory low-speed investigation of a 70 deg/46 deg cranked-wing planform was undertaken to evaluate two vortex-control concepts aimed at alleviating a severe pitch up which limits the usable lift well below the C(sub L,max) of the basic wing. One concept was a strake-like extension introduced across the wing crank, whose vortex helps to stabilize the outer-wing flow and alleviate tip stall. The other was a lower-surface cavity flap employed to trap a vortex just beneath the inboard leading edge, resulting in reduced vortex lift over the inner-wing panel. Each of these concepts was shown to eliminate the high-alpha pitch up, potentially raising the maximum usable lift of the cranked wing practically to its C(sub L,max) value

Subsonic investigations of vortex interaction control for enhanced high-alpha aerodynamics of a chine forebody/Delta wing configuration

A proposed concept to alleviate high alpha asymmetry and lateral/directional instability by decoupling of forebody and wing vortices was studied on a generic chine forebody/ 60 deg. delta configuration in the NASA Langley 7 by 10 foot High Speed Tunnel. The decoupling technique involved inboard leading edge flaps of varying span and deflection angle. Six component force/moment characteristics, surface pressure distributions and vapor-screen flow visualizations were acquired, on the basic wing-body configuration and with both single and twin vertical tails at M sub infinity = 0.1 and 0.4, and in the range alpha = 0 to 50 deg and beta = -10 to +10 degs. Results are presented which highlight the potential of vortex decoupling via leading edge flaps for enhanced high alpha lateral/directional characteristics

A low-speed wind tunnel study of vortex interaction control techniques on a chine-forebody/delta-wing configuration

A low speed wind tunnel evaluation was conducted of passive and active techniques proposed as a means to impede the interaction of forebody chine and delta wing vortices, when such interaction leads to undesirable aerodynamic characteristics particularly in the post stall regime. The passive method was based on physically disconnecting the chine/wing junction; the active technique employed deflection of inboard leading edge flaps. In either case, the intent was to forcibly shed the chine vortices before they encountered the downwash of wing vortices. Flow visualizations, wing pressures, and six component force/moment measurements confirmed the benefits of forced vortex de-coupling at post stall angles of attack and in sideslip, viz., alleviation of post stall zero beta asymmetry, lateral instability and twin tail buffet, with insignificant loss of maximum lift

A low speed wind tunnel investigation of Reynolds number effects on a 60-deg swept wing configuration with leading and trailing edge flaps

A low-speed wind tunnel test was performed to investigate Reynolds number effects on the aerodynamic characteristics of a supersonic cruise wing concept model with a 60-deg swept wing incorporating leading-edge and trailing-edge flap deflections. The Reynolds number ranged from 0.3 to 1.6 x 10 to the 6th, and corresponding Mach numbers from .05 to 0.3. The objective was to define a threshold Reynolds number above which the flap aerodynamics basically remained unchanged, and also to generate a data base useful for validating theoretical predictions for the Reynolds number effects on flap performance. This report documents the test procedures used and the basic data acquired in the investigation

The design and preliminary calibration of a boundary-layer flow channel

Design procedures for a new low-speed boundary-layer research channel are described. The channel is an open-circuit wind tunnel for the study of two-dimensional boundary layers under controlled pressure gradients, and follows design guidelines from published literature on blower tunnels with wide-angle diffusers. The contraction was arranged in a modular fashion permitting two different test sections of square and high-aspect-ratio cross section. A radical type of wide-angle diffuser was employed, and a stream-tube computer code (GE Streamtube Curvature Code) was used to check the contraction designs. The alternate test sections have the following specifications: 2- by 2-foot cross section with a fixed velocity of 23 ft/sec, and a boundary-layer section with a 0.5- by 2-foot cross section at a fixed velocity of approximately 89 ft/sec. Experimental techniques and data are described for the evaluation of diffuser effectiveness, boundary-layer channel characteristics, and overall performance of the facility

Impact of Fuselage Cross Section on the Stability of a Generic Fighter

Many traditional data bases, which involved smooth-sided forebodies, are no longer relevant for designing advanced aircraft. The current work provides data on the impact of chined-shaped fuselage cross section on the stability of a generic fighter configuration. Two different chined-shaped fuselages were tested upright and inverted. It was found that a fuselage with a 30° included chine angle resulted in significantly higher values of C L,max than a fuselage with a 100° included chine angle. This difference was attributed to the more beneficial vortical interaction between the stronger forebody vortices coming off of the sharper chine edges and the wing vortices. The longitudinal stability of the configuration with the sharper chine angle was also better because, based on pressures and flow visualization, the vortex burst over the wing was delayed until significantly higher values of a. Unstable rolling moment derivatives were also delayed to higher values of a for the sharper chine an..

Application of radial-splitters for improved wide-angle diffuser performance in blowdown tunnel

Severe flow separation in the 15:1 area-ratio, 38 ° total angle conical diffuse preceding the settling-chamber of an intermittent blowdown wind tunnel was eliminated by the use of a novel radial-splitter arrangement. As a consequence, the operating life of settling-chamber screens was greatly extended and test-section flow steadiness improved, with no penalty in the tunnel running time

Vortical flow management techniques

'Vortex management' refers to the purposeful manipulation and re-ordering of stable and concentrated vortical structures (e.g., resulting from flow separations from highly-swept leading edges and slender forebodies at moderate to high angles of attack) in order to enhance the aerodynamic performance and controllability of advanced, highly-maneuverable supersonic configurations. Exploratory experiments based on this approach have been conducted on generic research models at NASA Langley Research Center during recent years, investigating practical vortex flow control concepts and devices aimed at maneuver drag reduction, high angle of attack, pitch yaw and roll control, trimmed lift enhancement for short-field landing, etc. This paper reviews a selection of results attempting to clarify the basic aerodynamics of those concepts, and to evaluate their potential for improving performance and control. The vortex management concepts discussed herein include: aerodynamic compartmentation of highly-swept leading edges for alleviation of pitch non-linearities; capturing the leading edge vortex suction on forward-sloping flap surfaces for maneuver drag reduction; vortex lift modulation with articulated leading edge extensions for pitch-down and roll control at high angles of attack; vortex lift augmentation in the wing apex region to trim trailing edge flaps allowing shorter landing; and forebody vortex manipulation to alleviate uncontrolled asymmetry and also to generate yaw control in post-stall maneuvering. The precursor studies discussed here generally substantiated the vortex control concepts; questions such as configuration-sensitivity and scale effects are under continued investigation at NASA Langley and elsewhere