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At least 1,351 records · Page 75

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NASA NTRS1990

Electro-optical spin measurement system

An electro-optical spin measurement system for a spin model in a spin tunnel includes a radio controlled receiver/transmitter, targets located on the spin model, optical receivers mounted around the perimeter of the spin tunnel and the base of the spin tunnel for receiving data from the targets, and a control system for accumulating data from the radio controlled receiver and receivers. Six targets are employed. The spin model includes a fuselage, wings, nose, and tail. Two targets are located under the fuselage of the spin model at the nose tip and tail. Two targets are located on the side of the fuselage at the nose tip and tail, and a target is located under each wing tip. The targets under the fuselage at the nose tip and tail measure spin rate of the spin model, targets on the side of the fuselage at the nose tip and tail measure angle of attack of the spin model, and the targets under the wing tips measure roll angle of the spin model. Optical receivers are mounted at 90 degree increments around the periphery of the spin tunnel to determine angle of attack and roll angle measurements of the spin model. Optical receivers are also mounted at the base of the spin tunnel to define quadrant and position of the spin model and to determine the spin rate of the spin model.

Fodale, Robert↗

NASA NTRS1993

Investigation of high-alpha lateral-directional control power requirements for high-performance aircraft

Designers of the next-generation fighter and attack airplanes are faced with the requirements of good high angle-of-attack maneuverability as well as efficient high speed cruise capability with low radar cross section (RCS) characteristics. As a result, they are challenged with the task of making critical design trades to achieve the desired levels of maneuverability and performance. This task has highlighted the need for comprehensive, flight-validated lateral-directional control power design guidelines for high angles of attack. A joint NASA/U.S. Navy study has been initiated to address this need and to investigate the complex flight dynamics characteristics and controls requirements for high angle-of-attack lateral-directional maneuvering. A multi-year research program is underway which includes groundbased piloted simulation and flight validation. This paper will give a status update of this program that will include a program overview, description of test methodology and preliminary results.

Foster, John V.↗

NASA NTRS1993

Investigation of High-alpha Lateral-directional Control Power Requirements for High-performance Aircraft

Designers of the next-generation fighter and attack airplanes are faced with the requirements of good high-angle-of-attack maneuverability as well as efficient high speed cruise capability with low radar cross section (RCS) characteristics. As a result, they are challenged with the task of making critical design trades to achieve the desired levels of maneuverability and performance. This task has highlighted the need for comprehensive, flight-validated lateral-directional control power design guidelines for high angles of attack. A joint NASA/U.S. Navy study has been initiated to address this need and to investigate the complex flight dynamics characteristics and controls requirements for high-angle-of-attack lateral-directional maneuvering. A multi-year research program is underway which includes ground-based piloted simulation and flight validation. This paper will give a status update of this program that will include a program overview, description of test methodology and preliminary results.

Foster, John V.↗

NASA NTRS1990

Controlled vortical flow on delta wings through unsteady leading edge blowing

The vortical flow over a delta wing contributes an important part of the lift - the so called nonlinear lift. Controlling this vortical flow with its favorable influence would enhance aircraft maneuverability at high angle of attack. Several previous studies have shown that control of the vortical flow field is possible through the use of blowing jets. The present experimental research studies vortical flow control by applying a new blowing scheme to the rounded leading edge of a delta wing; this blowing scheme is called Tangential Leading Edge Blowing (TLEB). Vortical flow response both to steady blowing and to unsteady blowing is investigated. It is found that TLEB can redevelop stable, strong vortices even in the post-stall angle of attack regime. Analysis of the steady data shows that the effect of leading edge blowing can be interpreted as an effective change in angle of attack. The examination of the fundamental time scales for vortical flow re-organization after the application of blowing for different initial states of the flow field is studied. Different time scales for flow re-organization are shown to depend upon the effective angle of attack. A faster response time can be achieved at angles of attack beyond stall by a suitable choice of the initial blowing momentum strength. Consequently, TLEB shows the potential of controlling the vortical flow over a wide range of angles of attack; i.e., in both for pre-stall and post-stall conditions.

Lee, K. T.↗

NASA NTRS1961

Effects of Canard Planform and Wing-Leading-Edge Modification on Low-Speed Longitudinal Aerodynamic Characteristics of a Canard Airplane Configuration

An investigation has been conducted at low subsonic speeds to study the effects of canard planform and wing-leading-edge modification on the longitudinal aerodynamic characteristics of a general research canard airplane configuration. The basic wing of the model had a trapezoidal planform, an aspect ratio of 3.0, a taper ratio of 0.143, and an unswept 80-percent-chord line. Modifications to the wing included addition of full-span and partial-span leading-edge chord-extensions. Two canard planforms were employed in the study; one was a 60 deg sweptback delta planform and the other was a trapezoidal planform similar to that of the basic wing. Modifications to these canards included addition of a full-span leading-edge chord-extension to the trapezoidal planform and a fence to the delta planform. For the basic-wing-trapezoidal-canard configuration, rather abrupt increases in stability occurred at about 12 deg angle of attack. A slight pitch-up tendency occurred for the delta-canard configuration at approximately 8 deg angle of attack. A comparison of the longitudinal control effectiveness for the basic-wing-trapezoidal-canard combination and for the basic-wing-delta-canard combination indicates higher values of control effectiveness at law angles of attack for the trapezoidal canard. The control effectiveness for the delta-canard configuration, however, is seen to hold up for higher canard deflections and to higher angles of attack. Use of a full-span chord-extension deflected approximately 30 deg on the trapezoidal canard greatly improved the control characteristics of this configuration and enabled a sizeable increase in trim lift to be realized.

Spencer, Bernard, Jr.↗

NASA NTRS1993

Transitional flight characteristics of a geometrically simplified STOVL model

The transitional flight characteristics of a geometrically simplified Short Take-Off Vertical Landing (STOVL) aircraft configuration were measured in the NASA Ames 7- by 10-Foot Wind Tunnel. The experiment was designed to provide detailed data for evaluating the capability of computational fluid dynamics methods to predict the important powered lift flow parameters. The model consists of a 60 deg cropped delta wing planform; a blended fuselage; and tandem, circular, high pressure air jets that exit perpendicular to the flat lower surface. Freestream Mach number is limited to a maximum 0.2. Model angle of attack ranges from -10 deg to +20 deg. The nozzle pressure ratios of both jets are varied between 1 and 3, and the jet exit temperatures are maintained at near ambient conditions. Detailed surface pressure measurements show that suction pressure peaks located on the upper surface of the wing during conventional wingborne flight for angles of attack greater than 5 deg move to the wing lower surface at angles of attack less than 0 deg. A reduction in these suction pressure peaks is observed when the lift jets are operating. With sonic jet exit conditions, a freestream Mach number of 0.14, and 0 deg angle of attack, the jet-induced suckdown is equivalent to a 3.7 deg reduction in angle of attack. Schlieren, laser light sheet flow visualization and total pressure measurements in the jet plumes provide a description of the three-dimensional jet efflux flowfield.

Roth, Karlin R.↗

NASA NTRS1947

Investigation of a Spoiler-Type Lateral Control System on a Wing with Full-Span Flaps in the Langley 19-Foot Pressure Tunnel

Tests of a partial-span model of a large bomber-type air1ane were conducted to determine the. aerodynamic characteristics of the wing equipped with full-span flaps and a retractable spoiler end aileron lateral control system. The arrangement consisted of (1) a double slotted flap extending over aproximate1y 86 percent of the wing semispan, (2) a 20-percent constant-percentage-chord aileron extending from the outboard end of the flap to the wing tip, and (3) a retractable spoiler, located at the 65-percent wing-chord station and extending from approximately 63 percent of the wing semispan to the wing tip. In addition, tests were made of a wing vent (of 1 and 2 percent of the wing chord located directly behind the spoiler), perforations in the spoiler, a blot or cut-out along the lower edge of the spoiler and spoilers of various spans. With full-span flaps deflected and with the 2-percent vent open or closed the initial stalling of the wing occurred at the tips, but with the vents closed there probably would be no appreciable loss in lateral control until maximum lift was reached. The l-percent vent increased the rolling effectiveness of the spoiler at small spoi1er deflections, particularly at high angles of attack with flaps deflected. With flaps deflected the 2-percent vent caused a large reduction in both the wing lift and rolling effectiveness of the spoiler at large angles of attack. However, at small angle of attack the 2-percent vent increased the rolling effectiveness of the spoiler at small spoiler deflections. The simultaneous operation of the spoiler and vent (in contrast to a vent fixed in the wing) would result in a large increase in the effectiveness of the spoiler and would avoid any loss in wing lift as in a fixed vent arrangement. The tests of the spoiler modifications revealed that (1) the spoiler perforations reduced the rolling-moment and yawing-moment coefficients but caused the spoiler hinge-moment coefficients to become more positive; (2) the spoiler slot had no notable effect on the rolling-moment and yawing-moment characteristics but produced a positive increase in the spoiler hinge-moment coefficients at large spoiler deflections; (3) the effects produced by the individual modifications were additive when the various modifications were combined. In general, progressively decreasing the spoiler span by removing the segments from the inboard end of the spoiler caused a decrease in rolling effectiveness approximately proportional to the span of the segment.

WING FLAPS, SLOTTED & AILERONS↗

NASA NTRS1951

Preliminary Results Obtained from Flight Test of a 1/7-Scale Rocket-Powered Model of the Grumman XF10F Airplane Configuration in the Swept-Wing Condition, TED No. NACA DE 354

A flight investigation of a 1/7-scale rocket-powered model of the XF10F Grumman XFl0F airplane in the swept-wing configuration has been made. The purpose of this test was to determine the static longitudinal stability, damping in pitch, and longitudinal control effectiveness of the airplane with the center of gravity at 20 percent of the wing mean aerodynamic chord. Only a small amount of data was obtained from the test because, immediately after booster separation at a Mach number of 0.88, the configuration was directionally unstable and diverged in sideslip. Simultaneous with the sideslip divergence, the model became longitudinally unstable at 3 degree angle of attack and -6 degree sideslip and diverged in pitch to a high angle of attack. During the pitch-up the free-floating horizontal tail became unstable at 5 degree angle of attack and the tail drifted against its positive deflection limit.

Gardner, William N.↗

NASA NTRS1949

Wind-Tunnel Investigation of the Low-Speed Static Stability and Control Characteristics of a Model of Bell MX-776

An investigation has been made in the Langley stability tunnel to determine the low-speed static stability and control characteristics of a model of the Bell MX-776. The results of the investigation indicated that the basic model configuration was longitudinally stable in the angle-of-attack range from about -16 deg. to 16 deg. but that the stability was a minimum near O deg angle of attack. The data indicated an aerodynamic-center position about 0.64 body diameters behind the center of gravity at low angles of attack. Reduction in the size of the front horizontal fins increased the longitudinal stability. With 20 percent of the span of the normal front horizontal fins cut off the aerodynamic center was about 1.04 body diameters behind the center of gravity, and with front horizontal fins having the same area as the front vertical fins, the aerodynamic center was 2.26 body diameters behind the center of gravity (at low angles of attack).

Queijo, M. J.↗

NASA NTRS1983

A proposed simple and safe aircraft take-off or landing procedure with wing roughness or protuberances

A simple, safe takeoff and/or landing procedure has been developed for cases of wing roughness or proturbances. The procedure utilizes calculations of aerodynamic penalties and steady flight equations. Empirical relationships for lift, drag, and angle of attack penalties are used to construct the lift and drag coefficient versus the angle of attack. To maintain safe stall margin due to wing roughness, the lift coefficient and the angle of attack at takeoff or landing for a roughened aircraft must be less than that of a clean aircraft by the relative maximum lift loss and stall angle loss due to roughness. The procedure may be applied to many types of aircraft.

Dietenberger, M. A.↗

NASA NTRS1960

Experimental Determination of the Effects of Frequency and Amplitude of Oscillation on the Roll-Stability Derivatives for a 60 Degree Delta-Wing Airplane Model

A 60' delta-wing airplane model was oscillated in roll for several frequencies and amplitudes of oscillation to determine the effects of the oscillatory motion on the roll-stability derivatives for the model. The derivatives were measured at a Reynolds number of 1,600,000 for the wing alone, the wing-fuselage combination, and the complete model which included a triangular-plan-form vertical tail. Both rolling and yawing moments due to rolling velocity exhibited large frequency effects for angles of attack higher than 16 degrees. Variations in these derivatives were measured for the lowest frequencies of oscillation; as the frequency increased, the derivatives because more nearly linear with angle of attack. Both velocity derivatives were considerably different at high angles of attack from the corresponding derivatives measured by the steady-state rolling-flow technique. Rolling and yawing moments due to rolling acceleration were measured and similarly found to be highly dependent on frequency at high angles of attack. Some period and time-to-damp computations, which were made to reveal the significance of the acceleration derivatives, indicated that inclusion of the measured derivatives in the equations of motion lengthened the period of the lateral oscillation by 10 percent for a typical delta-wing airplane and increased the time to damp to one-half amplitude by 50 percent.

Fisher, Lewis R.↗

NASA NTRS2002

Three-Dimensional Effects on Multi-Element High Lift Computations

In an effort to discover the causes for disagreement between previous 2-D computations and nominally 2-D experiment for flow over the 3-clement McDonnell Douglas 30P-30N airfoil configuration at high lift, a combined experimental/CFD investigation is described. The experiment explores several different side-wall boundary layer control venting patterns, document's venting mass flow rates, and looks at corner surface flow patterns. The experimental angle of attack at maximum lift is found to be sensitive to the side wall venting pattern: a particular pattern increases the angle of attack at maximum lift by at least 2 deg. A significant amount of spanwise pressure variation is present at angles of attack near maximum lift. A CFD study using 3-D structured-grid computations, which includes the modeling of side-wall venting, is employed to investigate 3-D effects of the flow. Side-wall suction strength is found to affect the angle at which maximum lift is predicted. Maximum lift in the CFD is shown to be limited by the growth of all off-body corner flow vortex and consequent increase in spanwise pressure variation and decrease in circulation. The 3-D computations with and without wall venting predict similar trends to experiment at low angles of attack, but either stall too earl or else overpredict lift levels near maximum lift by as much as 5%. Unstructured-grid computations demonstrate that mounting brackets lower die the levels near maximum lift conditions.

Rumsey, Christopher L.↗

NASA NTRS1984

Inlet flow field investigation. Part 1: Transonic flow field survey

A wind tunnel investigation was conducted to determine the local inlet flow field characteristics of an advanced tactical supersonic cruise airplane. A data base for the development and validation of analytical codes directed at the analysis of inlet flow fields for advanced supersonic airplanes was established. Testing was conducted at the NASA-Langley 16-foot Transonic Tunnel at freestream Mach numbers of 0.6 to 1.20 and angles of attack from 0.0 to 10.0 degrees. Inlet flow field surveys were made at locations representative of wing (upper and lower surface) and forebody mounted inlet concepts. Results are presented in the form of local inlet flow field angle of attack, sideflow angle, and Mach number contours. Wing surface pressure distributions supplement the flow field data.

Yetter, J. A.↗

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 NTRS1980

Flow Visualization Study of the F-14 Fighter Aircraft Configuration

Water tunnel studies were performed to qualitatively define the flow field of the F-14. Particular emphasis was placed on defining the vortex flows generated at high angles of attack. The flow visualization tests were conducted in the Northrop water tunnel using a 1/72 scale model of the F-14 with a wing leading-edge sweep of 20 deg. Flow visualization photographs were obtained for angles of attack up to 55 deg and sideslip angles up to 10 deg. The F-14 model was investigated to determine the vortex flow field development, vortex path, and vortex breakdown characteristics as a function of angle of attack and sideslip. Vortex flows were found to develop on the highly swept glove and on the upper surface of the forebody. At 10 deg of sideslip, the windward glove vortex shifted inboard and broke down farther forward than the leeward glove vortex. This asymmetric breakdown of the vortices in sideslip contributes to a reduction in the lateral stability above 20 deg angle of attack. The initial loss of directional stability is a consequence of the adverse sidewash from the windward vortex and the reduced dynamic pressure at the vertical tails.

Lorincz, D. J.↗

NASA NTRSDate not supplied

Pterodactyl: 6-DOF Integration of Guidance and Control Algorithms in Genesis

The NASA-funded Pterodactyl project seeks to advance the current state-of-the-art for entry vehicles by developing novel guidance and control technologies for Deployable Entry Vehicles. This paper builds upon the Pterodactyl architecture that employed eight individually articulating flaps with two options for guidance, bank angle modulation with the Fully Numerical Predictor Corrector Entry Guidance (FNPEG) technique and angle of attack and sideslip modulation with FNPEG uncoupled range control. These, with a linear quadratic regulator controller, have previously been presented in separate 3-DOF trajectory simulations, one for translational motion and another for rotational dynamics. This work will show results for integrated 6-DOF simulations, leveraging recent advancements in trajectory simulation tools, namely the Julia-based Genesis flight simulation tool. Results show similar performance for both schemes when compared to the previously presented 3-DOF results once the controllers were adequately tuned. Angle of attack and sideslip angle modulation necessitated a controller designed at high dynamic pressure conditions, and bank angle modulation necessitated a controller designed at lower dynamic pressure conditions. Notably, the angle of attack and sideslip modulation approach could not achieve the parachute deploy point target Mach number of 2, when only tuning the controller gains, so future work will be needed to improve guidance and control tracking.

EDL↗

NASA NTRS1971

Hypersonic shock wave interaction and impingement

An experimental investigation was conducted on space shuttle type, body-wing configurations. The purpose of the investigation was to determine the effects of body and wing geometry on the hypersonic shock structure about these vehicles and on the resulting surface impingement of interior flow field shock and expansion waves. Schlieren photographs and thermographic phosphor paint data were obtained on three body cross sections with three wing planforms at 40, 50 and 60 degree angles of attack. Specific configuration data were obtained at 0 and 30 degree angles of attack to develop trends. These data were obtained at a nominal Mach number of 13.5 and a freestream unit Reynolds number of 0.7 million per foot. For comparison with these straight wing configurations, data were also obtained on a model of a point design, high cross-range, delta wing orbiter at 40, 50 and 60 degree angles of attack. As expected, the data on this delta wing orbiter indicated that the shock intersection/impingement phenomena associated with straight wing vehicles are considerably more complex than, and result in both windward and leeward surface heating regions not present on, the delta configuration.

Kessler, W. C.↗

NASA NTRS1995

Free-Flight Evaluation of Forebody Blowing for Yaw Control at High Angels of Attack

Forebody blowing is a concept developed to provide yaw control for aircraft flying at high angles of attack where a conventional rudder becomes ineffective. The basic concept is fairly simple. A small jet of air is forced out of the nose of the aircraft. This jet causes a repositioning of the forebody vortices in an asymmetrical fashion. The asymmetric forebody vortex flows develop a side force on the forebody which results in substantial yawing moments at high angles of attack. The purpose of this project was to demonstrate the use of forebody blowing as a control device through free-flight evaluation. This unique type of testing was performed at the NASA-Langley 30- by 60-foot tunnel. From these tests, it could then be shown that forebody blowing is an effective method of maintaining yaw control at high angles of attack.

Kiddy, Jason↗