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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 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Free-flight model investigation of a vertical-attitude VTOL fighter with twin vertical tails

Free-flight tests were conducted in the Langley full-scale tunnel to determine the stability and control characteristics of a vertical-attitude VTOL fighter having twin vertical tails and a pivoted fuselage forebody (nose-cockpit) arrangement. The flight tests included hovering flights and transition flights from hover to conventional forward flight. Static force tests were also made to aid in the analysis of the flight tests. The model exhibited satisfactory stability and control characteristics, and the transition from hovering flight to conventional forward flight was relatively smooth and straightforward.

Grafton, S. B.↗
NASA NTRS1975

Free-flight model investigation of a vertical-attitude VTOL fighter

Tests were made in the Langley full-scale tunnel and included a study of the stability and control characteristics of delta- and swept-wing configurations from hovering through the transition to normal forward flight. Static force tests were also conducted to aid in the analysis of the flight tests. With conventional artificial rate stabilization, very smooth transitions could be made consistently with relatively little difficulty. Because of the lower apparent damping and a tendency to diverge in yaw, however, the swept-wing configuration was considered to be much more difficult to fly than the delta-wing configuration. With rate dampers off, both configurations were very difficult to control and the control power needed for satisfactory flights was substantially higher than with the rate dampers operating.

Newsom, W. A., Jr.↗
NASA NTRS1975

Monte Carlo analysis of inaccuracies in estimated aircraft parameters caused by unmodeled flight instrumentation errors

An output error estimation algorithm was used to evaluate the effects of both static and dynamic instrumentation errors on the estimation of aircraft stability and control parameters. A Monte Carlo error analysis, using simulated cruise flight data, was performed for a high-performance military aircraft, a large commercial transport, and a small general aviation aircraft. The results indicate that unmodeled instrumentation errors can cause inaccuracies in the estimated parameters which are comparable to their nominal values. However, the corresponding perturbations to the estimated output response trajectories and characteristics equation pole locations appear to be relatively small. Control input errors and dynamic lags were found to be in the most significant of the error sources evaluated.

Hodge, W. F.↗
NASA NTRS1994

X-29 flight control system: Lessons learned

Two X-29A aircraft were flown at the NASA Dryden Flight Research Center over a period of eight years. The airplanes' unique features are the forward-swept wing, variable incidence close-coupled canard and highly relaxed longitudinal static stability (up to 35-percent negative static margin at subsonic conditions). This paper describes the primary flight control system and significant modifications made to this system, flight test techniques used during envelope expansion, and results for the low- and high-angle-of-attack programs. Through out the paper, lessons learned will be discussed to illustrate the problems associated with the implementation of complex flight control systems.

Clarke, Robert↗
NASA NTRS1980

Impact of longitudinal flying qualities upon the design of a transport with active controls

Direct constrained parameter optimization was used to optimally size a medium range transport for minimum direct operating cost. Several stability and control constraints were varied to study the sensitivity of the configuration to specifying the unaugmented flying qualities of transports designed with relaxed static stability. Additionally, a number of handling quality related design constants were studied with respect to their impact to the design.

Sliwa, S. M.↗
NASA NTRS1998

Including Aeroelastic Effects in the Calculation of X-33 Loads and Control Characteristics

Up until now, loads analyses of the X-33 RLV have been done at Marshall Space Flight Center (MSFC) using aerodynamic loads derived from CFD and wind tunnel models of a rigid vehicle. Control forces and moments are determined using a rigid vehicle trajectory analysis and the detailed control load distributions for achieving the desired control forces and moments, again on the rigid vehicle, are determined by Lockheed Martin Skunk Works. However, static aeroelastic effects upon the load distributions are not known. The static aeroelastic effects will generally redistribute external loads thereby affecting both the internal structural loads as well as the forces and moments generated by aerodynamic control surfaces. Therefore, predicted structural sizes as well as maneuvering requirements can be altered by consideration of static aeroelastic effects. The objective of the present work is the development of models and solutions for including static aeroelasticity in the calculation of X-33 loads and in the determination of stability and control derivatives.

Zeiler, Thomas A.↗
NASA NTRS1978

Flight evaluation of the transonic stability and control characteristics of an airplane incorporating a supercritical wing

A TF-8A airplane was equipped with a transport type supercritical wing and fuselage fairings to evaluate predicted performance improvements for cruise at transonic speeds. A comparison of aerodynamic derivatives extracted from flight and wind tunnel data showed that static longitudinal stability, effective dihedral, and aileron effectiveness, were higher than predicted. The static directional stability derivative was slower than predicted. The airplane's handling qualities were acceptable with the stability augmentation system on. The unaugmented airplane exhibited some adverse lateral directional characteristics that involved low Dutch roll damping and low roll control power at high angles of attack and roll control power that was greater than satisfactory for transport aircraft at cruise conditions. Longitudinally, the aircraft exhibited a mild pitchup tendency. Leading edge vortex generators delayed the onset of flow separation, moving the pitchup point to a higher lift coefficient and reducing its severity.

Matheny, N. W.↗
NASA NTRS1947

High-Speed Aerodynamic Characteristics of a 1/7-Scale Model of the Northrop YB-49 Airplane

Tests were conducted to find the effects of compressibility on the longitudinal stability and control of a 1/7-scale semispan model of the Northrop YB-49 airplane. Lift, drag, pitching moment, and elevon hinge moments were measured and are presented in graphical form. The results show that, due to a loss of lift on the outboard portion of the wing, the longitudinal static stability decreased rapidly as the Mach numbers increased above 0.735 the model experienced a climbing moment at positive lift coefficients. Also, a longitudinal-control effectiveness began to decrease at a Mach number of about 0.725

Robinson, Robert C.↗
NASA NTRS1959

Low-Subsonic Measurements of the Static and Oscillatory Lateral Stability Derivatives of a Sweptback-Wing Airplane Configuration at Angles of Attack from -10 to 90 deg

An investigation has been conducted in the Langley free-flight tunnel at low-subsonic speeds to provide some basic information on the stability and control characteristics in the high angle-of-attack range of an airplane configuration typical of current design trends. The investigation consisted of static- and dynamic-force tests over an angle-of- attack range from -10 to 90 deg. The dynamic-force tests, which consisted of both linear- and rotary-oscillation tests, were conducted at values of the reduced-frequency parameter k of 0.10, 0.15, and 0.20. The configuration was directionally unstable for all angles of attack above about 15 deg but maintained positive effective dihedral, control effectiveness, and damping in roll and yaw over most of the angle-of-attack range tested. The effects of frequency on the oscillatory stability derivatives were found to be generally small, but in a few cases the effects were relatively large.

Hewes, Donald E.↗
NASA NTRS1982

VTOL and VSTOL handling qualities specifications, an overview of the current status

The highlights of a comparative analysis between the current helicopter and VSTOL specifications and four representative rotary wing aircraft are presented. Longitudinal, lateral, and directional control power and dynamic stability characteristics were analyzed for hovering conditions. Forward flight static and dynamic stability were analyzed for the longitudinal and lateral-directional axes. Results of the analyses in terms of the applicability/utility of the MIL-H-8501A criteria are presented for each of the above areas. The review of the MIL-H-8301A criteria against those in MIL-F-83300 and AGARD 577 indicate many areas in which MIL-H-8501A does not give adequate design guidance.

Goldstein, K. W.↗
NASA NTRS1981

Thrust reversing effects on twin-engine aircraft having nonaxisymmetric nozzles

The effects of thrust reversing on stability, control and vertical tail loads of advanced fighter aircraft are reviewed. Several static test stand and wind tunnel investigations of nonaxisymmetric nozzles with integral thrust reversers are presented, and it is found that base drag on the aft face of reverser panels of blockers provides a significant contribution to in-flight thrust reverser performance. The location of tail surfaces relative to the thrust reverser has significant impact on control surface effectiveness, and depending on thrust reverser location, large vertical tail side force can result from the use of in-flight thrust reversers. It is concluded that highly effective nonaxisymmetric nozzle reversers can be designed to achieve a 50% static reverse thrust.

Capone, F. J.↗
NASA NTRS1984

An evaluation of the relative merits of wing-canard, wing-tail, and tailless arrangements for advanced fighter applications

Two sets of wind tunnel tests were performed to examine the relative merits of wing-canard, wing-tail and tailless configurations for advanced fighters. Both sessions focused on variable camber using automated, prescheduled leading and trailing edge flap positioning. The trials considered a modified F-16 tail and canard configuration at subsonic, transonic and supersonic speeds, a 60 deg delta wing sweep, a 44 deg leading edge trapezoidal wing at subsonic and supersonic speeds, vortex flow effects, and flow interactions in the canard-wing-tail-tailless variations. The results showed that large negative stabilities would need to be tolerated in wing-canard arrangements to make them competitive with wing-tail arrangements. Subsonic polar shapes for canard and tailless designs were more sensitive to static design margins than were wing-tail arrangements. Canards provided better stability at supersonic speeds. The static margin limits were a critical factor in control surface selection. Finally, a tailless delta wing configuration exhibited the lowest projected gross take-off weight and drag values.

Nicholas, W. U.↗
NASA NTRS1957

Flight Investigation of a Roll-stabilized Missile Configuration at Varying Angles of Attack at Mach Numbers Between 0.8 and 1.79

A missile research model was flown at supersonic speed to determine the quality of automatic roll stabilization at varying angles of attack. Aerodynamic rolling and pitching derivatives were determined from the flight record. It was concluded that the combination of the gyro-actuated automatic pilot with wing-tip ailerons provided adequate roll stabilization under conditions encountered in flight.

INSTRUMENTS↗
NASA NTRS1950

Static Longitudinal Stability of a Tandem-Coupled Bomber-Fighter Airplane Configuration Similar to One Proposed by Douglas Aircraft Company, Inc.

At the request of the Air Materiel Command, an investigation was made in the Langley free-flight tunnel to determine the longitudinal stability and control characteristics of models coupled together in a tandem configuration for aerial refueling similar to one proposed by the Douglas Aircraft Company, Inc. Static force tests were made with 1/20-scale models of the B-29 and F-80 airplanes to determine the effects of rigidly coupling the airplanes together. The Douglas configuration differs from the rigid configuration tested in that it provides for some freedom in pitch and vertical displacement. The force tests showed that, for the bomber alone, the aerodynamic center was 0.21 mean aerodynamic chord behind the center of gravity (stable) but that for the tandem configuration with rigid coupling the aerodynamic center was 0.28 mean aerodynamic chord forward of the center of gravity of the combination (unstable). This reduction in stability was caused by the downwash of the bomber on the fighter. The pitching moment produced by elevator deflection of the bomber was reduced approximately 50 percent by addition of the fighter. Some recent flight tests made in the free-flight tunnel on models in a similar tandem configuration indicated that, with a hinged coupling permitting freedom in pitch, the stability of the combination was better than that obtained with a rigid coupling and was about the same as that for the bomber alone.

Hewes, Donald E.↗
NASA NTRS1972

Static aerodynamic characteristics of a model with a 17 percent thick supercritical wing

The statis stability and control characteristics and air loads on the wing and aileron of a 0.09 scale model of an airplane with a 17 percent thick supercritical wing were investigated. The longitudinal aerodynamic characteristics were determined over a lift coefficient range sufficiently high to induce buffet onset at Mach numbers from 0.30 to 0.80. In addition, the lateral aerodynamic characteristics were determined as a function of sideslip angle and angle of attack at selected Mach numbers.

Ferris, J. C.↗
NASA NTRS1992

Pitch Control Margin at High Angle of Attack - Quantitative Requirements (Flight Test Correlation With Simulation Predictions)

Aircraft designs that employ relaxed static stability (RSS) have the following problem: reduced pitching moments associated with RSS at high angle of attack (AOA) require a minimum pitch recovery moment or margin to guarantee a safe return from high AOA maneuvers at the most aft center of gravity (CG) encountered during a mission. Recent incidents and mishaps on Class IV aircraft have demonstrated a need for establishing quantitative longitudinal high AOA pitch control margin design guidelines for future aircraft. The Naval Air Warfare Center - Aircraft Division (NA WC-AD) is currently supporting an effort in conjunction with NASA Langley Research Center (NASA LaRC) to quantify such requirements. NASA LaRC has conducted a series of extensive simulation evaluations to define these design guidelines. The purpose of flight tests were to validate the overall research test methodology by comparing pilot comments, pilot ratings, and aircraft response characteristics gathered during inflight recoveries from high AOA conditions to those gathered during the fixed-base simulation sessions. Tests were completed on an F/A-18A in six flights for a total of 9.8 flight hours using an AOA and CG buildup sequence. Flight test results have validated the simulation studies in that pilot rating of high AOA nose-down recoveries were based on the short-term response interval in the forms of pitch acceleration and rate. In addition, flight test has demonstrated that high AOA pitch control margin can be evaluated using a stabilized pushover method.

Flight Testing↗