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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1961

The Effect of Lateral-Directional Control Coupling on Pilot Control of an Airplane as Determined in Flight and in a Fixed-Base Flight Simulator

A flight and fixed-base simulator study was made of the effects of aileron-induced yaw on pilot opinion of aircraft lateral-directional controllability characteristics. A wide range of adverse and favorable aileron-induced yaw was investigated in flight at several levels of Dutch-roll damping. The flight results indicated that the optimum values of aileron- induced yaw differed only slightly from zero for Dutch-roll damping from satisfactory to marginally controllable levels. It was also shown that each range of values of aileron-induced yawing moment considered satisfactory, acceptable, or controllable increased with an increase in the Dutch- roll damping. The increase was most marked for marginally controllable configurations exhibiting favorable aileron-induced yaw. Comparison of fixed-base flight simulator results with flight results showed agreement, indicating that absence of kinesthetic motion cues did not markedly affect the pilots' evaluation of the type of control problem considered in this study. The results of the flight study were recast in terms of several parameters which were considered to have an important effect on pilot opinion of lateral-directional handling qualities, including the effects of control coupling. Results of brief tests with a three-axis side-arm controller indicated that for control coupling problems associated with highly favorable yaw and cross-control techniques, use of the three-axis controller resulted in a deterioration of control relative to results obtained with the conventional center stick and rudder pedals.

Vomaske, Richard F.↗
NASA NTRS1929

Contribution to the aileron theory

In an attempt to treat theoretically the effect of ailerons, difficulty arises because an aileron may begin at any point of the wing. Hence the question arises as to how the transition of the lift distribution proceeds at such a point, since the effect of the aileron (i.e., the moment generated about the longitudinal axis) depends largely on this distribution. In order to answer this question regarding the lift distribution during irregular variations in the angle of attack at first independently of other influences, especially those of the wing tips, we have taken as the basis of the following theoretical discussion a wing of infinite span and constant chord which exhibits at one point an irregular variation in the angle of attack. As regards the mathematical treatment, we will first consider a wing with periodically recurring angle of attack.

LIFT DISTRIBUTION - WINGS↗
NASA NTRS1950

Transonic Stability and Control Investigation of a 1/80-Scale Model of the Consolidated Vultee Skate 9 Seaplane, TED No. NACA DE 345: Transonic-Bump Method

An investigation of the longitudinal stability and of the all-movable horizontal tail and aileron control of a 1/80-scale reflection-plane model of the Consolidated Vultee Skate 9 seaplane has been made through a Mach number range of 0.6 to 1.16 on the transonic bump of the Langley high-speed 7- by 10-foot tunnel. At moderate lift coefficients (0.4 to 0.8) and below a Mach number of 1.0 the model was statically unstable longitudinally. The static longitudinal stability of the model at low lift coefficients increased with Mach number corresponding to a shift in aerodynamic center from 37 percent mean aerodynamic chord at a Mach number of 0.60 to 64 percent at a Mach number of 1.10. Estimates indicate that the tail deflection angle required for steady flight and for accelerated maneuvers of the Skate 9 airplane would probably not vary greatly with Mach number at sea level, but for accelerated maneuvers at altitude the tail deflection angle would probably vary erratically with Mach number. The variation of rolling-moment coefficient with aileron deflection angle was approximately linear, agreed well with theory, and held for the range of aileron deflections tested (-17.1 deg to 16.6 deg). At low lift coefficients the drag rise occurred at Mach numbers of 0.95 and 0.90 for the wing alone and the complete model, respectively. The effects of the canopy on the model were small. For the ranges investigated, angle-of-attack and Mach number changes caused no large pressure drops in the jet-engine duct.

Riebe, John M.↗
NASA NTRS1976

Theoretical Stability and Control Characteristics of Wings With Various Amounts of Taper and Twist

Stability derivatives have been computed for twisted wings of different planforms that include variations in both the wing taper and the aspect ratio. Taper ratios of 1.0, 0.50, and 0.25 are considered for each of three aspect ratios: 6, 10, and 16. The specific derivatives for which results are given are the rolling moment and the yawing moment derivatives with respect to rolling velocity, yawing velocity, and angle of sideslip. In addition to the stability derivatives, results are included for determining the theoretical rolling moment due to aileron deflection and a series of influence lines is given by which the loading across the span may be determined for any angle-of-attack distribution that may occur on the wing planforms considered.

Henry A Pearson↗
NASA NTRS1979

Nonlinear decoupled control synthesis for maneuvering aircraft

A control law for decoupling roll rate, angle of attack and sideslip in rapid, nonlinear airplane maneuvers is derived. For simplicity, only moments caused by ailerons, rudder and elevator are considered, and control forces are neglected. Simulated responses of the closed loop system, including the control forces, show that the neglected forces have no significant effects and that large, simultaneous lateral and longitudinal maneuvers can be precisely performed. The violent divergences which occur in open-loop maneuvers of this type are eliminated by the decoupling. The control law showed little sensitivity to 20 percent perturbation in three important stability derivatives.

Singh, S. N.↗
NASA NTRS1938

Theoretical stability and control characteristics of wings with various amounts of taper and twist

Stability derivatives have been computed for twisted wings of different plan forms that include variations in both the wing taper and the aspect ratio. Taper ratios of 1.0, 0,50, and 0.25 are considered for each of three aspect ratios: 6, 10, and 16. The specific derivatives for which results are given are the rolling-moment and the yawing-moment derivatives with respect to (a) rolling velocity, (b) yawing velocity, and (c) angle of sideslip. These results are given in such a form that the effect of any initial symmetrical wing twist (such as may be produced by flaps) on the derivatives may easily be taken into account. In addition to the stability derivatives, results are included for determining the theoretical rolling moment due to aileron deflection and a series of influence lines is given by which the loading across the span may be determined for any angle-of-attack distribution that may occur on the wing plan forms considered. The report also includes incidental references to the application of the results.

Pearson, Henry A↗
NASA NTRS1947

Investigation of the Stability and Control Characteristics of a 1/20-Scale Model of the Consolidated Vultee XB-53 Airplane in the Langley Free-Flight Tunnel

An investigation of the low-speed, power-off stability and control characteristics of a 1/20-scale model of the Consolidated Vultee XB-53 airplane has been conducted in the Langley free-flight tunnel. In the investigation it was found that with flaps neutral satisfactory flight behavior at low speeds was obtainable with an increase in height of the vertical tail and with the inboard slats opened. In the flap-down slat-open condition the longitudinal stability was satisfactory, but it was impossible to obtain satisfactory lateral-flight characteristics even with the increase in height of the vertical tail because of the negative effective dihedral, low directional stability, and large-adverse yawing moments of the ailerons.

Bennett, Charles V.↗
NASA NTRS1953

Rocket-Model Measurements of Zero-Lift Damping in Roll of the Bell MX-776 Missile at Mach Numbers from 0.6 to 1.56

The zero-lift damping in roll of the Bell MX-776 missile has been measured by a sting-mounted rocket-model technique at Mach numbers from 0.6 to 1.56. The damping-in-roll data, in general, show no unusual variation with Mach number. Aileron rolling-moment effectiveness derived from these data and previously obtained rolling-effectiveness data appear reasonable,

Bland, William N., Jr.↗
NASA NTRS1937

Full-scale Wind-tunnel and Flight Tests of a Fairchild 22 Airplane Equipped with External-airfoil Flaps

Wind-tunnel and flight tests have been made of a Fairchild 22 airplane equipped with a wing having external-airfoil flaps that also perform the function of ailerons. Lift, drag, and pitching-moment coefficients of the airplane with several flap settings, and the rolling- and yawing-moment coefficients with the flaps deflected as ailerons were measured in the full-scale tunnel with the horizontal tail surfaces and propeller removed. The effect of the flaps on the low speed and on the take-off and landing characteristics, the effectiveness of flaps when used as ailerons, and the forces required to operate them as ailerons were determined in flight. The wind-tunnel tests showed that the flaps increased the maximum lift coefficient of the airplane from 1.51 with the flap in the minimum drag position to 2.12 with the flap in the minimum drag position to 2.12 with the flap deflected 30 degrees. In the flight tests the minimum speed decreased from 46.8 miles per hour with the flaps up to 41.3 miles per hour with the flaps deflected. The required take-off run to attain a height of 50 feet was reduced from 820 to 750 feet and the landing run from a height of 50 feet was reduced from 930 to 480 feet. The flaps for this installation gave lateral control that was not entirely satisfactory. Their rolling action was good but the adverse yaw resulting from their use was greater than is considerable, and the stick forces required to operate them increased too rapidly with speed.

Reed, Warren D↗
NASA NTRS1994

Supersonic aerodynamic characteristics of a circular body Earth-to-Orbit vehicle

The circular body configuration is a generic single- or multi-stage reusable Earth-to-orbit transport. A thick clipped-delta wing is the major lifting surface. For directional control, three different vertical fin arrangements were investigated: a conventional aft-mounted center fin, wingtip fins, and a nose-mounted fin. The tests were conducted in the Langley Unitary Plan Wind Tunnel. The configuration is longitudinally stable about the estimated center of gravity of 0.72 body length up to a Mach number of about 3.0. Above Mach 3.0, the model is longitudinally unstable at low angles of attack but has a stable secondary trim point at angles of attack above 30 deg. The model has sufficient pitch control authority with elevator and body flap to produce stable trim over the test range. The model with the center fin is directionally stable at low angles of attack up to a Mach number of 3.90. The rudder-like surfaces on the tip fins and the all-movable nose fin are designed as active controls to produce artificial directional stability and are effective in producing yawing moment. The wing trailing-edge aileron surfaces are effective in producing rolling moment, but they also produce large adverse yawing moment.

Ware, George M.↗
NASA NTRS1928

The Effect of a Flap and Ailerons on the N.A.C.A. M-6 Airfoil Section

This report contains the results obtained at the Langley Memorial Aeronautical Laboratory on an N. A. C. A. M-6 airfoil, fitted with a flap and ailerons, and tested in the variable density wind tunnel at a density of 20 atmospheres. Airfoil characteristics are given for the model up to 48 degree angle of attack with the flap set at various angles, and also with the ailerons set at similar angles. The approximate lift distribution and the center of pressure variation along the span are determined with the model at 18 degree angle of attack and with the ailerons displaced at 20 degrees. Approximate rolling moment and yawing moment coefficients are determined for the various aileron settings. A comparison of the calculated angles of zero lift and the calculated lift and moment coefficiencies with those observed is given in the appendix.

Higgins, George J↗
NASA NTRS1945

Comparison of Pitching Moments Produced by Plain Flaps and by Spoilers and Some Aerodynamic Characteristics of an NACA 23012 Airfoil with Various Types of Aileron

Sectional characteristics of airfoil having retractable slotted flap with plain, slot-lip, or retractable ailerons are presented for a large range of aileron deflections. The analysis indicated that pitching moments produced by spoilers were less positive than those produced by plain flaps of equal effectiveness, also that pitching moments created by the spoiler increased less with the Mach number than similar moments produced by plain flaps. Positive values of pitching moment decreased as devices were located nearer airfoil leading edge.

PITCHING MOMENTS - AIRFOILS↗
NASA NTRS2015

Novel Control Effectors for Truss Braced Wing

At cruise flight conditions very high aspect ratio/low sweep truss braced wings (TBW) may be subject to design requirements that distinguish them from more highly swept cantilevered wings. High aspect ratio, short chord length and relative thinness of the airfoil sections all contribute to relatively low wing torsional stiffness. This may lead to aeroelastic issues such as aileron reversal and low flutter margins. In order to counteract these issues, high aspect ratio/low sweep wings may need to carry additional high speed control effectors to operate when outboard ailerons are in reversal and/or must carry additional structural weight to enhance torsional stiffness. The novel control effector evaluated in this study is a variable sweep raked wing tip with an aileron control surface. Forward sweep of the tip allows the aileron to align closely with the torsional axis of the wing and operate in a conventional fashion. Aft sweep of the tip creates a large moment arm from the aileron to the wing torsional axis greatly enhancing aileron reversal. The novelty comes from using this enhanced and controllable aileron reversal effect to provide roll control authority by acting as a servo tab and providing roll control through intentional twist of the wing. In this case the reduced torsional stiffness of the wing becomes an advantage to be exploited. The study results show that the novel control effector concept does provide roll control as described, but only for a restricted class of TBW aircraft configurations. For the configuration studied (long range, dual aisle, Mach 0.85 cruise) the novel control effector provides significant benefits including up to 12% reduction in fuel burn.

White, Edward V.↗
NASA NTRS1954

Method for calculating the rolling and yawing moments due to rolling for unswept wings with or without flaps or ailerons by use of nonlinear section lift data

The methods of NACA reports 865 and 1090 have been applied to the calculation of the rolling-moment and yawing-moment coefficients due to rolling for unswept wings with or without flaps or ailerons. The methods are based on lifting-line theory and allow the use of nonlinear section lift data. The method presented in this report permits calculations to be made somewhat beyond maximum lift for wings having no twist or continuous twist and employing airfoil sections which do not display large discontinuities in the lift curves. Calculations can be made up to maximum lift for wings with discontinuous twist such as that produced by partial-span flaps or ailerons, or both. Two calculated examples are presented in simplified computing forms in order to illustrate the procedures involve.

Martina, Albert P↗
NASA NTRS1953

Method for calculating the rolling and yawing moments due to rolling for unswept wings with or without flaps or ailerons by use of nonlinear section lift data

The methods of NACA Reports 865 and 1090 have been applied to the calculation of the rolling- and yawing-moment coefficients due to rolling for unswept wings with or without flaps or ailerons. The methods allow the use of nonlinear section lift data together with lifting-line theory. Two calculated examples are presented in simplified computing forms in order to illustrate the procedures involved.

CONTROLS - COMPLETE WINGS↗
NASA NTRS1933

The Effects of Slots and Flaps on Lateral Control of a Low-wing Monoplane as Determined in Flight

This paper presents the results of flight tests made to determine the effect of slots and flaps on the lateral control of a low-wing monoplane. Maximum angular accelerations in roll and yaw produced by sudden application of the ailerons and maximum accelerations in yaw produced by sudden application of the rudder during gliding flight were recorded for the following wing arrangements: (a) no auxiliary device; (b) full-span slots; (c) plain flaps; (d) flaps and full-span slots; (e) wing-tip slots. Rolling- and yawing-moment coefficients were derived from the accelerations. The full span slots and the flaps each had about the same influence on the aileron rolling moments. At values of the lift coefficient obtainable with the plain wing, the effect of these devices was negligible. At the higher lift coefficients obtainable with these devices, the rolling-moment coefficients increased slightly but, despite this increase, the aileron effectiveness progressively decreased with increasing lift coefficient, owing to the corresponding reduction in air speed. In the range covered by the tests, the effectiveness of the controls was appreciably reduced by the wing-tip slots. The adverse yawing moment of the ailerons experienced at the large lift coefficients obtained with the flaps was appreciably less than at similar lift coefficients obtained with either full-span or wing-tip slots. The yawing moments produced by the rudder were only slightly affected by the use of the auxiliary devices. The airplane was found to be laterally unstable with all combinations tested. Because of the angular velocities acquired in the time taken to deflect the ailerons the rolling moments recorded in flight were only about two thirds the values that would have been obtained with the wing restrained as in wind-tunnel tests.

WING SLOTS↗