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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 595 records · Page 33
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Lee surface flow phenomena over Space Shuttle at large angles of attack at M infinity = 6

An investigation was conducted regarding the surface heat transfer on the leeward side of the Space Shuttle, taking into account different free stream Reynolds numbers and angles of attack. Another investigation was concerned with the peak heating due to boundary layer transition and flow separation in the case of different Space Shuttle configurations. Criteria for vortex generation on the leeward flow were also examined along with the effect of transition on vortex generation. Attention was given to the leeward flowfield in the vortex flow region and the method of constructing an equivalent model for leeward flowfield analysis.

Zakkay, V.↗
NASA NTRS1980

Sensor for measuring instantaneous angle of attack of helicopter blades

Systematic investigations were performed on a variety of probes to determine their potential for possible application as sensors attached to helicopter blades to measure both the instantaneous angle of attack as well as the dynamic head during actual flight operations. After some preliminary considerations a sensor of essentially spherical shape, about 30 mm in diameter, was designed. The sensor was provided with three pressure ports, and it housed two pressure transducers required for sensing the prevailing pressures acting outside on the surface. The sensors were subsequently tested in the laboratory under a variety of flow conditions to determine their aerodynamic characteristics. Two series of tests were performed: in the first series the sensor was fixed in space while exposed to steady uniform flow, while in the second series the sensor was made to oscillate, thus simulating the cyclic pitch change of the helicopter blades. While the cyclic pitch frequencies were of about the same magnitude as encountered in flight, the flow velocities during tests fell well below those experienced in a rotating blade. The tests showed that the sensors performed satisfactorily under low subsonic flow conditions with frequencies not exceeding five Hz.

Barna, P. S.↗
NASA NTRS1984

Flight determination of the aerodynamic stability and control characteristics of the NASA SGS 1-36 sailplane in the conventional and deep stall angles-of-attack of between -5 and 75 degrees

The flight test procedure and the preliminary analysis of the results obtained from twenty manned flights of the SGS 1-36 in the high angles of attack Deep Stall region are discussed. A comparison of the flight determined stability and control derivatives, those of the wind tunnel, and the estimated aerodynamic data is also presented. Furthermore, deep stall dynamics response of the SGS 1-36 is discussed briefly to explain some of the unexpected flight observations.

Mahdavi, F. A.↗
NASA NTRS1989

In-flight flow visualization characteristics of the NASA F-18 high alpha research vehicle at high angles of attack

Surface and off-surface flow visualization techniques have been used to visualize the three-dimensional separated flows on the NASA F-18 high alpha research vehicle at high angles of attack. Results near alpha = 25 deg to 26 deg and alpha = 45 deg to 49 deg are presented. Both the forebody and leading-edge extension (LEX) vortex cores and breakdown locations were visualized using smoke. Forebody and LEX vortex separation lines on the surface were defined using an emitted-fluid technique. A laminar separation bubble was also detected on the nose cone using the emitted fluid technique and was similar to that observed in the wind-tunnel test, but not as extensive. Regions of attached, separated, and vortical flow were noted on the wing and the leading-edge flap using tufts and flow cones, and compared well with limited wind-tunnel results.

Fisher, David F.↗
NASA NTRS1936

Wind-Tunnel Research Comparing Lateral Control Devices Particularly at High Angles of Attack XIII : Auxiliary Airfoils Used as External Ailerons

This is the thirteenth report on a series of systematic tests comparing lateral control devices with particular reference to their effectiveness at high angles of attack. The present wind tunnel tests were made to determine the most feasible locations for lateral control surfaces mounted externally to a rectangular Clark y wing.

Weick, Fred E↗
NASA NTRS1994

Parameter identification for X-31A at high angles of attack

The U.S./German experimental aircraft X-31A was designed and constructed to demonstrate enhanced fighter maneuverability. Post-stall maneuvering is enabled by applying new technologies such as high angle of attack aerodynamics and flight control system integrated thrust vectoring. Two demonstrator aircraft have been built by the main contractors, Rockwell International and Deutsche Aerospace (formerly MBB). Flight testing started in October 1990 and before the end of 1992 both aircraft had accomplished a significant number of flights covering the entire AoA regime from about -5 to 70 deg. Throughout the envelope expansion, DLR Institute of Flight Mechanics conducted parameter identification (PID) to determine the aerodynamic parameters of the aircraft from flight test data and to compare the results to the predictions from the aerodynamic dataset (ADS). The application of system identification to high AoA / post-stall flight data raises some major problems, which are discussed in this paper. Results from both longitudinal and lateral-directional motion will be presented.

S Weiss↗
NASA NTRS2003

Prediction of Flows about Forebodies at High-Angle-of-Attack Dynamic Conditions

A Reynolds-average Navier Stokes method developed for rotorcraft type of flow problems is applied for predicting the forces and moments of forebody models at high-angle-of-attack dynamic conditions and for providing insight into the flow characteristics at these conditions. Wind-tunnel results from rotary testing on generic forebody models conducted by NASA Langley and DERA are used for comparison. This paper focuses on the steady-state flow problem.

vanDam, C. P.↗
NASA NTRS1980

Exploratory investigation of the effects of vortex bursting on the high angle-of-attack lateral-directional stability characteristics of highly-swept wings

A recent low-speed wind-tunnel investigation of highly-swept wings has shown that vortex breakdown at high angles of attack can cause large destabilizing effects on static lateral-directional stability characteristics, and that the destabilizing effects of vortex breakdown can be greatly aggravated by model support strut interference effects. The present paper discusses these effects based on the results of static force tests of several highly-swept wing configurations for different wind-tunnel strut setup arrangements. Also included in the paper are photographs obtained during tuft-, smoke-, and helium-bubble flow visualization studies to indicate wing flow behavior patterns

Johnson, J. L., Jr.↗
NASA NTRS1982

Convergence behavior that controls adaptive wind tunnel walls near the test section in the high angle of attack range

The NACA 0012 profile at Mach 0.5 was investigated in a wind tunnel with adaptive walls. It is found that adaptation of the flexible walls is possible in the high angle of attack range on both sides of maximum lift. Oil film photographs of the flow at the profile surface show three dimensional effects in the region of the corners between the profile and the sidewall. It is concluded that pure two dimensional separated flow is not possible.

Ziemann, J.↗
NASA NTRS1983

Computational technique for three-dimensional compressible flow past wings at high angles of attack

A computational technique, based on the integral solution of the full potential equation, has been developed for the solution of three-dimensional high-subsonic flows past wings at high angles of attack. The problem includes two sources of nonlinearities: a boundary oriented nonlinearity (separated flow roll up) and a region oriented nonlinearity (flow compressibility). The former is represented by a nonlinear vortex lattice while the latter is represented by a source distribution inside a finite volume. The solution is obtained by using double iteration cycles; a separated flow (wake) iteration cycle and a compressibility iteration cycle. The computational technique is applied to a delta wing. The results show that the technique is very promising and efficient.

Kandil, O. A.↗
NASA NTRS2005

A Tail Buffet Loads Prediction Method for Aircraft at High Angles of Attack

Aircraft designers commit significant resources to the design of aircraft in meeting performance goals. Despite fulfilling traditional design requirements, many fighter aircraft have encountered buffet loads when demonstrating their high angle-of-attack maneuver capabilities. As a result, during test or initial production phases of fighter development programs, many new designs are impacted, usually in a detrimental way, by resulting in reassessing designs or limiting full mission capability. These troublesome experiences usually stem from overlooking or completely ignoring the effects of buffet during the design phase of aircraft. Perhaps additional requirements are necessary that addresses effects of buffet in achieving best aircraft performance in fulfilling mission goals. This paper describes a reliable, fairly simple, but quite general buffet loads analysis method to use in the initial design phases of fighter-aircraft development. The method is very similar to the random gust load analysis that is now commonly available in a commercial code, which this analysis capability is based, with some key modifications. The paper describes the theory and the implementation of the methodology. The method is demonstrated on a JSF prototype example problem. The demonstration also serves as a validation of the method, since, in the paper, the analysis is shown to nearly match the flight data. In addition, the paper demonstrates how the analysis method can be used to assess candidate design concepts in determining a satisfactory final aircraft configuration.

Pototzky, Anthony S.↗
NASA NTRS1976

Investigation of two bifurcated-duct inlet systems from Mach 0 to 2.0 over a wide range of angles of attack

A 15.354 percent/scale lightweight fighter type inlet/forebody was tested over a Mach number range of 0 to 2.0. Model configurations consisted of side mounted normal shock and fixed overhead ramp type inlets. Each configuration consisted of two inlets ducted (bifurcated) to supply a single engine face. The normal shock inlet variables included a boundary layer splitter bleed system, alternate boundary layer splitter plates, alternate upper and lower cowl lip shapes, and a blow-in-door (auxiliary inlet) in one lower lip. The only variable of the fixed overhead ramp inlet was the boundary layer bleed flow. Reynolds numbers ranged from 7.6 x 1 million to 19.5 x 1 million/m. Angle of attack ranged from -10 to 35 deg and angle of sideslip from -8 to 8 deg. Test measurements included engine face total pressure recovery, steady state distortion, dynamic distortion, and surface static pressures on the forebody and inlet surfaces.

Latham, E. A.↗
NASA NTRS1957

Impingement of Cloud Droplets on 36.5-Percent-Thick Joukowski Airfoil at Zero Angle of Attack and Discussion of Use as Cloud Measuring Instrument in Dye-Tracer Technique

The trajectories of droplets i n the air flowing past a 36.5-percent-thick Joukowski airfoil at zero angle of attack were determined. The amount of water i n droplet form impinging on the airfoil, the area of droplet impingement, and the rate of droplet impingement per unit area on the airfoil surface were calculated from the trajectories and cover a large range of flight and atmospheric conditions. With the detailed impingement information available, the 36.5-percent-thick Joukowski airfoil can serve the dual purpose of use as the principal element in instruments for making measurements in clouds and of a basic shape for estimating impingement on a thick streamlined body. Methods and examples are presented for illustrating some limitations when the airfoil is used as the principal element in the dye-tracer technique.

Brun, R. J.↗
NASA NTRS1984

Airfoil computation at high angles of attack, inviscid and viscous phenomena

An implicit central difference code is used to calculate two dimensional inviscid and thin-layer Navier-Stokes solutions for flow about an NACA0012 airfoil at high angles of attack. Among the issues addressed are whether separation can occur in an inviscid calculation and what the causes would be of such separation. Examples are shown of inviscid shocked flow with and without separation and shock-free flow with separation. An Euler solution with self-induced oscillation and separation driven by a strong shock is contrasted with a shock-free solution whose separation is caused by numerical error. Computed solutions to the Euler equations are compared to those of the potential equations. Comparisons are also made between experimental data from wind tunnel tests and viscous calculations at similar conditions.

Barton, J. T.↗
NASA NTRS1933

Wind-tunnel research comparing lateral control devices, particularly at high angles of attack IV : floating tip ailerons on rectangular wings

This report is the fourth of a series on systematic tests conducted which compares lateral control devices with particular reference to their effectiveness at high angles of attack. The present report covers tests with floating tip ailerons on rectangular Clark y wings. Ailerons of two profiles were tested - symmetrical and Clark y, both with adjustable trailing-edge flaps. Each form was tested at three hinge-axis locations, both with and without vertical end plates between the ailerons and the wing proper. The results from these tests are compared with the results from tests on a wing of the same over-all size equipped with average-sized ordinary ailerons.

Weick, Fred E↗
NASA NTRS1961

Static Longitudinal Aerodynamic Characteristics at Transonic Speeds of a Blunted Right Triangular Pyramidal Lifting Reentry Configuration for Angles of Attack up to 110 deg

The static longitudinal aerodynamic characteristics of a blunted rig ht triangular pyramidal lifting reentry configuration have been obtai ned at Mach numbers from 0.60 t o 1.20 and angles of attack from 0 de grees to 110 degrees. The test Reynolds number, based on the mean geometric chord, varied during the tests from about 2.2 x 10 (exp 6) to 5 .7 x 10 (exp 6).

Mugler, John P., Jr.↗
NASA NTRS1974

Low speed aerodynamic characteristics of an 0.075-scale F-15 airplane model at high angles of attack and sideslip

An 0.075 scale model representative of the F-15 airplane was tested in the Ames 12 foot pressure wind tunnel at a Mach number of 0.16 to determine static longitudinal and lateral directional characteristics at spin attitudes for Reynolds numbers from 1.48 to 16.4 million per meter (0.45 to 5.0 million per foot). Angles of attack ranged from 0 to +90 deg and from -40 deg to -80 deg while angles of sideslip were varied from -20 deg to +30 deg. Data were obtained for nacelle inlet ramp angles of 0 to 11 deg with the left and right stabilators deflected 0, -25 deg, and differentially 5 deg and -5 deg. The normal pointed nose and two alternate nose shapes were also tested along with several configurations of external stores. Analysis of the results indicate that at higher Reynolds numbers there is a slightly greater tendency to spin inverted than at lower Reynolds numbers. Use of a hemispherical nose in place of the normal pointed nose provided an over correction in simulating yawing moment effects at high Reynolds numbers.

Petroff, D. N.↗