Experimental and Computational Investigation of Spoiler Deployment on Wing Stall

Upper surface flaps commonly referred to as spoilers or drag brakes can increase maximum lift, and improve aerodynamic efficiency at high, near-stall angles of attack. This phenomenon was studied experimentally and computationally using a 0.307626 m chord length NACA 2412 airfoil in six different configurations, and one baseline clean configuration. A wind tunnel model was placed in the Ryerson Low Speed Wind Tunnel (atmospheric, closed-circuit, 3 ft × 3 ft test section) at a Reynold’s number of approximately 780,000 and a Mach number of 0.136. The wind tunnel study increased the lift coefficient by 0.393%-2.497% depending on the spoiler configuration. A spoiler of 10% chord length increased the maximum lift coefficient by 2.497 % when deflected 8º, by 2.110% when deflected 15º, and reduced the maximum lift coefficient by 2.783% when deflected 25º. A spoiler of 15% chord length produced smaller maximum lift coefficient gains; 0.393% when deflected 8º, by 1.760% when deflected 15º, and reduced the maximum lift coefficient by 4.475% when deflected 25º. Deflecting the spoiler increased the stall angle between 37.658% and 87.544% when compared with the clean configuration. The drag coefficient of spoiler configurations was lower than the clean configuration at angles of attack above 18º. The combination of the increased lift and reduced drag at angles of attack above 18º created by the spoiler configurations resulted in a higher aerodynamic efficiency than the clean configuration case. A 10% chord length spoiler deflected at 8º produced the highest aerodynamic efficiency gains. At low angles of attack, the computational study produced consistently higher lift coefficients compared with the wind tunnel experiment. The lift-slope was consistent with the wind tunnel experiment lift-slope. The spoiler airfoil stall behaviour was inconsistent with the results from the wind tunnel experiment. The drag coefficient results were consistent with the wind tunnel experiment at low angles of attack. However, the spoiler equipped airfoils did not reduce drag at high angles of attack. Therefore, the computational model was not valid for the spoiler configurations at high angles of attack.

Experimental and Computational Investigation of Spoiler Deployment on Wing Stall

Upper surface flaps commonly referred to as spoilers or drag brakes can increase maximum lift, and improve aerodynamic efficiency at high, near-stall angles of attack. This phenomenon was studied experimentally and computationally using a 0.307626 m chord length NACA 2412 airfoil in six different configurations, and one baseline clean configuration. A wind tunnel model was placed in the Ryerson Low Speed Wind Tunnel (atmospheric, closed-circuit, 3 ft × 3 ft test section) at a Reynold’s number of approximately 780,000 and a Mach number of 0.136. The wind tunnel study increased the lift coefficient by 0.393%-2.497% depending on the spoiler configuration. A spoiler of 10% chord length increased the maximum lift coefficient by 2.497 % when deflected 8º, by 2.110% when deflected 15º, and reduced the maximum lift coefficient by 2.783% when deflected 25º. A spoiler of 15% chord length produced smaller maximum lift coefficient gains; 0.393% when deflected 8º, by 1.760% when deflected 15º, and reduced the maximum lift coefficient by 4.475% when deflected 25º. Deflecting the spoiler increased the stall angle between 37.658% and 87.544% when compared with the clean configuration. The drag coefficient of spoiler configurations was lower than the clean configuration at angles of attack above 18º. The combination of the increased lift and reduced drag at angles of attack above 18º created by the spoiler configurations resulted in a higher aerodynamic efficiency than the clean configuration case. A 10% chord length spoiler deflected at 8º produced the highest aerodynamic efficiency gains. At low angles of attack, the computational study produced consistently higher lift coefficients compared with the wind tunnel experiment. The lift-slope was consistent with the wind tunnel experiment lift-slope. The spoiler airfoil stall behaviour was inconsistent with the results from the wind tunnel experiment. The drag coefficient results were consistent with the wind tunnel experiment at low angles of attack. However, the spoiler equipped airfoils did not reduce drag at high angles of attack. Therefore, the computational model was not valid for the spoiler configurations at high angles of attack.

Pitching effect on transonic wing stall of a blended flying wing with low aspect ratio

Numerical simulation of the pitching effect on transonic wing stall of a blended flying wing with low aspect ratio was performed using improved delayed detached eddy simulation (IDDES). To capture the discontinuity caused by shock wave, a second-order upwind scheme with Roe’s flux-difference splitting is introduced into the inviscid flux. The artificial dissipation is also turned off in the region where the upwind scheme is applied. To reveal the pitching effect, the implicit approximate-factorization method with sub-iterations and second-order temporal accuracy is employed to avoid the time integration of the unsteady Navier–Stokes equations solved by finite volume method at Arbitrary Lagrange–Euler (ALE) form. The leading edge vortex (LEV) development and LEV circulation of pitch-up wings at a free-stream Mach number M = 0.9 and a Reynolds number Re = [Formula: see text] is studied. The Q-criterion is used to capture the LEV structure from shear layer. The result shows that a shock wave/vortex interaction is responsible for the vortex breakdown which eventually causes the wing stall. The balance of the vortex strength and axial flow, and the shock strength, is examined to provide an explanation of the sensitivity of the breakdown location. Pitching motion has great influence on shock wave and shock wave/vortex interactions, which can significantly affect the vortex breakdown behavior and wing stall onset of low aspect ratio blended flying wing.

Ground Effect on the Vortex Flow and Aerodynamics of a Slender Delta Wing

The ground effect on the aerodynamic loading and leading-edge vortex (LEV) flow generated by a slender delta wing was investigated experimentally. Both the lift and drag forces were found to increase with reducing ground distance (up to 50% of the wing chord). The lift increment was also found to be the greatest at low angles of attack α and decreased rapidly with increasing ground distance and α. The ground effect-caused earlier wing stall was also accompanied by a strengthened LEV with an increased rotational speed and size compared to the baseline wing. The smaller the ground distance, the stronger the LEV and the earlier vortex breakdown became. Meanwhile, the vortex trajectory was also found to be located further inboard and above the delta wing in ground effect compared to its baseline-wing counterpart. Finally, for wing-in-ground effect (WIG) craft with delta-wing planform the most effective in-ground-effect flight should be kept within 10% of the wing chord.

A-10 Wing Leading Edge Effects on Engine Stability: Part 1 — Analysis and Evaluation of Wing Leading Edge Configurations

The A-10 aircraft has fuselage mounted engines with inlets just above the rear of the wing. The A-10 employs a deployable slat system to delay wing stall directly in front of the engines. Wing stall can lead to high inlet distortion and ultimately engine stall for this aircraft. To enhance overall performance of the A-10 Close Air Support Aircraft, wing leading-edge designs that do not employ slats were considered. Fifteen potential wing leading-edge proposals including drooped wings, wings with fences, wings with vortex generators, an optimized slat and a specially designed wing were evaluated through test and analysis for replacing the A-10 slat system. The performance of the wing leading-edge candidates were characterized by their inlet engine distortion effect on the loss of stability pressure ratio (ΔPRS) on the TF-34 engine fan and compressor. The drooped wings or “droops” were designated by the amount of droop in a percent of chord. Droops tested were 3, 5, 7, 10, and 10-5% twisted (5% outboard, 10% inboard). The 7, 10, and 10-5% droops were tested with outboard fences. The 10% droop and designed wing were tested with vortex generators. The paper discusses the previous work and technical basis for selecting the wing leading edge candidates, the analysis tools and techniques, the test and analysis of the candidate configurations, the overall effectiveness of the best candidate.

A-10 Wing Leading Edge Effects on Engine Stability: Part 2 — Identifying Total Pressure Distortion Sources and the Effects on Engine/Airframe Integration

The A-10 aircraft has fuselage mounted engines with inlets just above the rear of the wing. A deployable slat system is used on the A-10 to delay wing stall directly in front of the engines. Wing stall can lead to high inlet distortion and ultimately engine stall for this aircraft. Many alternate wing leading-edge designs were recently considered for replacement of the slat system to reduce maintenance cost, reduce system complexity and increase system reliability. Fifteen potential wing leading-edge proposals for replacing the A-10 slat system were evaluated through test and analysis. Performance of the wing leading-edge candidates was characterized by the effect of inlet/engine distortion on loss of stability pressure ratio (ΔPRS) for the TF-34 engine fan and compressor. The many protuberances and non-aerodynamic shapes of the A-10 outer mold lines can generate flow structures that cause significant inlet/engine total pressure distortion. Thorough understanding of these flow structures and their impact on inlet/engine distortion was necessary to fully assess the performance of candidate wing leading edge configurations. The paper discusses the integrated test and evaluation tools and methods used to identify the sources of inlet/engine total pressure distortion and the associated impact to engine/airframe integration.