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By D.M. Rao Principal investigator: G.L. Goglis Final Report
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For the period June 22, 1979 Jute 21, 1960 Prepared for the national Aeronautics and Space Administration Langley Research Center Aampton. Virginia Under Cooperative Agreement nCCI-9.
John S. Peterson, Jr., Technical Monitor Subsonic-Transonic Aerodynamics Division .q
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September 1961 DP,PARTMENT OF MECHANICAL ENGINEERING AND MECHANICS SCHOOL OF ENGINEERING OLD DOMINION UNIVERSITY NORFOLK, VIRGINIA CARGO MIND-TUNNEL STUDIES OF ADVANCED AIRCRAFT CONCEPTS By D. M. Rao Principal Investigator: G.L. Goglia Final Report For the period June 22, 1979 - June 21, 1980 Prepared for the National Aeronautics and Space Administration Langley Research Center Hampton, Virginia 23665 Under Cooperative Agreement NCCI-9
B. Peterson, Jr., Technical Monitor
John Subsonic-Transonic Aerodynamics Division Submitted by the Old Dominion University Research Foundation P.O. Box 6369
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Norfolk, Virginia 23508-0369 September 1981 WIND-TUNNEL STUDIES OF ADVANCED CARGO AIRCRAFT CONCEPTS By D.M. Rao* INTRODUCTION The principal accomplishments of this research effort are summarized as f ol.lows: (1) Cats analysis of wind-:unnel tests on novel leading —edge devices wing model at higher angles of for drag-reduction on a 60- degree delta attack was completed. A data report (NASA CR-159120) and an ALAA paper (No.
80-0310) highlighting the main results were prepared. This research was submitted as a Master's thesis by Mr. T.D. Johnson, Jr.
(2) 4 preliminary wind-tunnel test of the l-eading-elge vortex-flap concept on a 14-degree :'.elta wing was documented (NASA CR-159161). Consid- erable interest has been generated within the aerospace industry in the potential of this device for improving the subsonic performance of super- sonic cruise airplanes. A wind-tunnel test program was conducted by General Dynamics at NASA/Langley Research Center (LaRC) to evaluate vortex flaps on thair fighter model. Vortex flaps were designed for a Boeing Recce-Strike model for testing at NASA/LaRC in March 1980.
Results of subsonic wind-tunnel tests on a NASA arrow-wing super- (3) sonic transport configuration equipped with segmented vortex flaps were presented at the Langlev Supersonic Cruise Research Conference (SCR-79).
The L/D improvements at lift coefficients corresponding to climb and landing-approach by the use of vorte flaps were confirmed during a follow- up test in the NASA/LaRC V/STOL tunnel with a larger scale model.
(4) A paper summarizing the vortex flap research was accepted for presentation at the International Council of Aeronautical Sciences (ICAS) Conference in Munich in October 1980.
* Research Professor, Department of Mechanical Engineering and Mechanics, Old Dominion University, Norfolk, Virginia 23508.
(5) Two new technology reports hove also been prepared, and these are detailed in the following sections.
NEW TECHNOLOGY REPORT NO. 1 Description of the Problem Highly avt a pt slender wings employed on supersonic-cruise aircraft and the subsonic phases of flight missiles are characteristically inefficient in at high angles of attack (as in approach, climb, and combat maneuver). flow separation from the leading edges generates vortices which provide addi- tional lift but at a high cost in drag, resulting in poor lift/drag ratio.
Ala , the center of vortex lift moves towards the nose, causing pitch-up difficulties, particularly when vortex breakdown is encountered. The vortex benavior also influences lateral/directional stability and often itac-feres unfavorably with tail controls. Conventional means to delay separation ani the onset of vortices, e.g., by deflecting the entire leading edge down in order to maintain attached flow to a higher angle of attack, are limited in 'rhe large practice bu mechanical complexity and attendant weight penalty.
deflection angles required raise the li'.elihood of flow separation inboard along the hingaline so that complete vortex suppression on the wing may not be possible uy drooping the lead edges.
Description of the New Technology The basis of the present aproach is not to prevent the natural tendency of flow separation and vortex formation on higher swept wings, but rather to utilize the vortex in s different way. Separation is aliowed to occur on a highly deflected leading-edge flap and the resulting vortex held on the flap upper surface all across the span (see fig. 1). The vortex auction acting on the flap area then produces an aerodynamic thrust component which is responsible for drag reduction.
The optimum flow pattern conceived for the "vortex flap" requires flow attachment to occur just at the wing-flap junction, as indicated in figure 1. Thuc+, not only is the entire flap area brought under the influence of the vortex, but also the entry to the wing is smooth, resulting in attacl flow conditions on the wing. This optimum flow, whi:h can be approached suitably adjusting the flap angle for any given angle of attack, can be exactly obtained with a planar flap only at one spanwise position due to Oe three-dimensional nature of the wing flow field However, by means of a segmented vortex flap, where each segment may be independently adjusted, the optimum Clow condition may be closely approximated over a large portion of the span. Moreover, some degree of pitching moment adjustment for longi- tudinal trimming may be availaule by appropriate setting of the fore and aft segments, while still enjoying the drag-reduction benefit of the vortex flap principle in a good measure. Segmented flaps, as illustrated in figure 2, may also be found more practical from actuation considerations on large aircraft than a single-piece flap.
Unique Features of the Technology and Results of its Application A unique feature of the vortox flap is that, throughout the angle of attack range, the flow type remains qualitatively unchanged. Accordingly, no large or sudden change in the aerodynamic characteristics, as happens when forcibly maintained attached flow suddenly reverts to separation, will occur with the vortex flap.
Typical wind-tunnel teat data are presented in figure 3 to show the drag-reduction potential of the vortex flap concept applied to a supersonic cruise airplane configuration. The segmented vortex flap data are compared with an ideal leading-edge camber shape determined experimentally for fully attached flow. This ideal shape cannot be regarded as a practical variable- geometry leading-edge structure; by contrast, the vortex flaps are simple panels shaped to retract into the lower surface of the wing when not used.
The vortex ;laps produce the same order of lift/drag ratio improvements as the ideally cambered attached-flow leading edge over a Lange of lift coeffi- cients appropriate to subsonic flight. In addition, they also alleviated significantly the excessive dihedral effect of the basic wing and so improved the cross-wind landing characteristics.
Comments The vortex flap concept was proposed and initial results proving the principle obtained before a similar idea called th• "vortex tab" was announced by Boeing in 1478. While Boeing has reported some flow visual- ization experiments to indicate the formation of the flap vortex, no data indicating the performance of the device have yet been presented. The vortex flap results measured at NASA/LaRC by Dr. D.M. Rao are the firs' to definitively establish the effectiveness and potential of the technology.
NEW TECHNOLOGY REPORT NO. 2 Description of the Problem One of the limitations on subsonic maneuverability of highly swept, slender wing combat aircraft arises from degraded effectiveness of conven- tional ailerons and spoilers employed for roll control. These control surfaces, which ::re essentially dependent for efficient operation on attached flow over the wing upper surface, are rendered largely ineffective when large-scale separation occurs at the leading edges at high angles of attack. The loss of roll controllability in the high-lift condition, when lateral stability and roll-damping are also at reduced levels, progressively leads to determination of tracking ability, handling difficulties, and lowered resistance to departure.
Description of the New Technology The new technology offers a means of aerodynamically augmenting the roll-control effectiveness on slender wings at higher angles of attack by manipulating the vortex flow generated from leading-edge separation.
The proposed manipulator takes the form of a flap hinged at or close to the leading edge, normally retracted flush with the wing upper suface to ^--form to the airfoil shape. In operation, the flap on either wing panel s raised to a suitehle angle (see fig. 4). The flow patterns resulting from deployment of the flap are essentially of two type-j (A and B) depending on the angle of attack, as depicted in figure 5. Considering first the relevant high angle-of-attack case B, a system of two vortices is obtained, one on the flap upper surface and the other inboard on the wing. 1be redistribution of spanwise-lift caused by the vortex system B, alsc, shown schematically In figure 5, mo ,^es the center-of-lift towards the wing panel with deployed flap and produces the desired rolling moment.
At low angles of attack the flow pattern A (where the first vortex occurs below the leading edge) is likaly to produce an cpposite rolling moment. While the low angle-of-attack case is not relevant to the roll- control function, it suggests another use for the upper-Aurface vortex flaps (viz as airbrakes) which will be described later.
A cypicol set of subsoaic wind-tunnel test data (obtained in HA5A/1.aRC 7 x 10 ft high-speed tunnel) for an upper-surface vortex flap simulated on the left panel of a 74-degree delta wing research model is presented in f igure 6.
The flap drag at low angles of attack (flow pattern A, fig. 5) is manifested as a positive side-force component which reduces to zero and then reverses sign as the flow pattern changes to B at a higher angle of attack and the suction of the first vortex begins to be felt on the flap. This also marks the origin of a positive rolling moment, increasing linearly with lift coeffizient (up to C L - 0.8 for the particular flap configuration), which is accompanied by an acceptable yewing moment.
As evident from the drag polar comparia- . a shown in figure 5, flap deployment for roll control incurs no measurable drag penalty. It may be inferred that the thrust component of the flap force fully compensated the skin-friction drag of the flap surface.
Conventional elevons will still be needed for roll-control in the cruise phase of flight, and a combination of elevons and upper - surface vortex flap may therefore be considered for use at high angles of attack.
Wind-tunnel data presented is figure 7 shows that for the test configuration the two contributions to rolling moment were essentially additive. However,
a synergistic effect i^ indicated in the C L
range from U.b to 1. 0, which may be attributed to the suction induced by the inboard vortex as it passes over the down-deflected eleven.
In an alternate mode of operation, the upper surface vortex Elays on both wing panels may be deployed together. This typically affects the lift/urag characteristics as indicated in figure 8. A large drag is incurred at low _ift coefficients (< 0.2) associated with the flow patters A, which my be utilized for deceleration from high-soeed flight. Pitching- trim change in moment measurements (fig. 8) indicate that there will be n ,)
symmetri ^ flap deployment at high speed. In the range C L -
0.4 to O. b, the flow pattern B on the upper surface produces L/D improvements on the order of 10 percent. The magnitude of these effects is dependent on the flap size and deflection angle.
C Unique Featur6s of the Technology and Results of its Application The ,.nique and advantageous features of the upper surface vortex flap as a roll-control device, as illustrated by preliminary wind-tunnel tests on u conceptual model briefly discussed under "Description of the New Tech- nology," may be summarized as follows: (1) It is not based on attached-flow requirement as are the conven- tional control. surfaces, V% , : operates by manipulating three-dimensional flow separation and vortices naturally present on highly swept slender wings at large angles of attack.
(2) Rolling moment increases linearly with lift coefficient up to high
values of lift (CL -
1.0).
(3) Induced yawing moment is acceptable: i.e., it acts in tiie direc- tion of the tern, therefore req«iring little rudder assistance.
(4) No drag penalty is incurred due to control deployment.
° (5) It augments roll power synergistically when used in combination with conventional elevons.
In addition, by deploying the surface vortex flaps symmetrically on both wing panels: (6) A powerful airbrake effect is available for rapid and controlled deceleration from high speed with no trim cnangs. The bteady nature of the vortex flow field is expected to yield exceptionally low buffet level during deceleration and to avoid wake-excitation of empennage structures.
(7) Aerodynamic thrust effect from the vortex-induced suction on the flaps improves the lift/drag ratio by 5 to 10 percent in low-speed flight (CL - 0.4 to 0.6).
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Figure 1. Vortex flap concept.
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Figure 4. Upper-surface vortex flap concept for roll control.
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