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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1978

Canard-body-tail missile test at angles of attack to 50 deg in the Ames 11-foot transonic wind tunnel

Blunted ogive cylinder missile models with a length-to-diameter ratio of 10.4 were tested at transonic speeds and large angles of attack in an 11 foot transonic wind tunnel. The configurations are: body, body with tail panels, body with canards, and body with canards and tails. Forces and moments from the entire model and each of the eight fins were measured over the pitch range of 20 deg to 50 deg, and roll angles of 0 deg to 45 deg and canard deflection angles between 0 deg and 15 deg. The Reynolds number ranged from 3.9 x 10 to the 6th power per meter. Large side forces and yawing moments were observed for some of the test cases involving a symmetric geometry.

Allen, C. Q.↗
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↗
NASA NTRS1936

Spinning Characteristics of Wings I : Rectangular Clark Y Monoplane Wing

A series of wind tunnel tests of a rectangular Clark Y wing was made with the NACA spinning balance as part of a general program of research on airplane spinning. All six components of the aerodynamic force and moment were measured throughout the range of angles of attack, angles of sideslip, and values omega b/2v likely to be attained by a spinning airplane; the results were reduced to coefficient form. It is concluded that a conventional monoplane with a rectangular Clark y wing can be made to attain spinning equilibrium throughout a wide range of angles of attack but that provision of a yawing moment coefficient of -0.02 (against the spin) by the tail, fuselage, and interferences will insure against attainment of equilibrium in a steady spin.

Bamber, M J↗
NASA NTRS1974

Effects of several factors on theoretical predictions of airplane spin characteristics

The influence of different mathematical and aerodynamic models on computed spin motion was investigated along with the importance of some of the aerodynamic and nonaerodynamic quantities defined in these models. An analytical technique was used which included the aerodynamic forces and moments acting on a spinning aircraft due to steady rotational flow and the contribution of the rotary derivatives to the oscillatory component of the total angular rates. It was shown that (1) during experimental-analytical correlation studies, the flight-recorded control time histories must be faithfully duplicated since the spinning motion can be sensitive to a small change in the application of the spin entry controls; (2) an error in the assumed inertias, yawing moments at high angle of attack, and initial spin entry bank angle do not influence the developed spin significantly; (3) damping in pitch derivatives and the center of gravity location play a role in the spinning motion; and (4) the experimental spin investigations conducted in a constant atmospheric density environment duplicate the Froude number only at the initial full-scale spin altitude (since the full-scale airplane at high altitudes experiences large density changes during the spin.)

Bihrle, W., Jr.↗
NASA NTRS1994

Vertical Axis Rotational Motion Cues in Hovering Flight Simulation

A previous study that examined how yaw motion affected a pilot's ability to perform realistic hovering flight tasks indicated that any amount of pure yaw motion had little-to-no effect on pilot performance or opinion. In that experiment, pilots were located at the vehicle's center of rotation; thus lateral or longitudinal accelerations were absent. The purpose of the new study described here was to investigate further these unanticipated results for additional flight tasks, but with the introduction of linear accelerations associated with yaw rotations when the pilot is not at the center of rotation. The question of whether a yaw motion degree-of-freedom is necessary or not is important to government regulators who specify what simulator motions are necessary according to prescribed levels of simulator sophistication. Currently, specifies two levels of motion sophistication for flight simulators: full 6-degree-of-freedom and 3-degree-of-freedom. For the less sophisticated simulator, the assumed three degrees of freedom are pitch, roll, and heave. If other degrees of freedom are selected, which are different f rom these three, they must be qualified on a case-by-case basis. Picking the assumed three axes is reasonable and based upon experience, but little empirical data are available to support the selection of critical axes. Thus, the research described here is aimed at answering this question. The yaw and lateral degrees of freedom were selected to be examined first, and maneuvers were defined to uncouple these motions from changes in the gravity vector with respect to the pilot. This approach simplifies the problem to be examined. For this experiment, the NASA Ames Vertical Motion Simulator was used in a comprehensive investigation. The math model was an AH-64 Apache in hover, which was identified from flight test data and had previously been validated by several AH-64 pilots. The pilot's head was located 4.5 ft in front of the vehicle center of gravity, which is representative of the AH-64 pilot location. Six test pilots flew three tasks that were specifically designed to represent a broad class of situations in which both lateral and yaw motion cues may be useful. For the first task, the pilot controlled only the yaw axis and was required to rapidly acquire a North heading from 15 deg yaw offsets to either the East or West. This task allowed for full, or 1:1, motion to be used in all axes (yaw, lateral, and longitudinal). The second task was a 10 sec., 180 deg. pedal turn over a runway, but with the pilot only controlling the yaw degree-of-freedom. The position of the vehicle's center-of-mass remained fixed. This maneuver was taken from a current U.S. Army rotary wing design standard5 and is representative of a maneuver performed for acceptance of military helicopters; however, it does not allow for full 1:1 motion, since the simulator cab cannot rotate 180 deg. The third task required the pilot to perform a rapid 9 ft climb at a constant heading. This task was challenging, because rapid collective lever movement in the unaugmented AH64 results in a substantial yawing moment (due to engine torque) that must be countered by the pilot. This task also had full motion in all axes, but, in this case, the pilot had two axes to control simultaneously, rather than one as in the previous tasks. Four motion configurations were examined for each task: full motion (except for the 180 deg turn, for which the motion system was configured to provide as much motion as possible), full linear with no yaw motion, full yaw with no linear motion, and no motion. Each configuration was flown four times in a randomized test matrix, and the pilots were not informed of the configuration given. Vehicle state data were recorded for objective performance comparisons, and pilots provided subjective comments and ratings. As part of the pilots' evaluation, they were asked to rate the compensation required, the overall fidelity of the motion as compared to real flight, and whether motion was detected or not in each of the six degrees of freedom. In addition, the pilots provided a numerical level-of confidence rating, between 1 and 7, corresponding to how sure they were whether or not motion was present in each degree-of-freedom. The latter rating allow classical signal detection analysis to be performed.

Schroeder, Jeffrey A.↗
NASA NTRS1979

Aerodynamic interactions from reaction controls for lateral control of the M2-F2 lifting-body entry configuration at transonic and supersonic and supersonic Mach numbers

Tests were conducted in the Ames 6 by 6 foot wind tunnel to determine the interaction of reaction jets for roll control on the M2-F2 lifting-body entry vehicle. Moment interactions are presented for a Mach number range of 0.6 to 1.7, a Reynolds number range of 1.2 x 10 to the 6th power to 1.6 x 10 to the 6th power (based on model reference length), an angle-of-attack range of -9 deg to 20 deg, and an angle-of-sideslip range of -6 deg to 6 deg at an angle of attack of 6 deg. The reaction jets produce roll control with small adverse yawing moment, which can be offset by horizontal thrust component of canted jets.

Bailey, R. O.↗
NASA NTRS2011

Analysis of Asymmetric Aircraft Aerodynamics Due to an Experimental Wing Glove

Aerodynamic analysis on a business jet with a wing glove attached to one wing is presented and discussed. If a wing glove is placed over a portion of one wing, there will be asymmetries in the aircraft as well as overall changes in the forces and moments acting on the aircraft. These changes, referred to as deltas, need to be determined and quantified to make sure the wing glove does not have a drastic effect on the aircraft flight characteristics. TRANAIR, a non-linear full potential solver was used to analyze a full aircraft, with and without a glove, at a variety of flight conditions and angles of attack and sideslip. Changes in the aircraft lift, drag and side force, along with roll, pitch and yawing moment are presented. Span lift and moment distributions are also presented for a more detailed look at the effects of the glove on the aircraft. Aerodynamic flow phenomena due to the addition of the glove and its fairing are discussed. Results show that the glove used here does not present a drastic change in forces and moments on the aircraft, but an added torsional moment around the quarter-chord of the wing may be a cause for some structural concerns.

Hartshorn, Fletcher↗
NASA NTRS1982

Control of the forebody vortex orientation by asymmetric air injection. Part A: Application to enhance departure/spin recovery of fighter aircraft. Part B: Details of the flow structure

A novel concept which was developed to provide powerful directional control effectiveness for a fighter aircraft at high angles of attack, where more traditional controls have very limited capability is discussed. The concept utilizes the energy concentrated in the strong forebody vortices (which form on slender bodies at high relative incidence) by controlling the lateral orientation of the vortices with respect to the body. The present concept seeks to utilize the inherent sensitivity of the vortex positioning and its bistable nature to an advantage allowing control of the forces which are developed. As it turns out, the direction or sense of the asymmetric vortex pair is much easier to control than to attenuate. The work which was done to develop the concept for application to an aircraft is described and is directed toward the effects of the concept on aircraft forces and moments and on the flight mechanics of the aircraft during maneuvering at high angles of attack. The objective was to utilize the side force associated with asymmetric vortices, in a controlled manner, to enhance the ability of the fighter to recover from a departure from controlled flight. The results from these water tunnel and wind tunnel experiments show that a small amount of tangential blowing along the forebody near the apex can effectively alter the forebody vortex system and generate large restoring yawing moments.

Skow, A. M.↗
NASA NTRS1957

Aerodynamic Loads on Tails at High Angles of Attack and Sideslip

Results are presented for the loads and moments acting on the individual tail surfaces of a body-tail combination over a wide range of angles of attack and sideslip. The effects of forebody length and panel-panel interference on the characteristics are included. It is shown that large nonlinear variations in these loads and moments, which occur at some combinations of angle of attack and sideslip, cannot be predicted by low-angle theory. A relatively simple, but general, theoretical method for calculating these load and moment characteristics is described, and the results from this method are found to be in good agreement with experiment provided the initial positions of the forebody vortices are known. It is shown that a simple application of slender-body theory can be used to predict the side loads due to sideslip that are contributed by a vertical tail on a wide variety of wing-body-tail combinations at low angles of attack. For several configurations, changes are indicated which reduced the vertical-tail loads per unit yawing moment of each complete configuration at large angles of attack. Some results are presented on the effect of high angle of attack on the induced-flow field and tail loads due to a wing at supersonic speed.

SLENDER BODY↗
NASA NTRS1934

The Aerodynamic Forces and Moments Exerted on a Spinning Model of the NY-1 Airplane as Measured by the Spinning Balance

A preliminary investigation of the effects of changes in the elevator and rudder settings and of small changes in attitude upon the aerodynamic forces and moments exerted upon a spinning airplane was undertaken with the spinning balance in the 5-foot vertical tunnel of the National Advisory Committee for Aeronautics. The tests were made on a 1/12-scale model of the ny-1 airplane. Data by which to fix the attitude, the radius of spin, and the rotational and air velocities were taken from recorded spins of the full-scale airplane. Two spinning conditions were investigated. All six components of the aerodynamic reaction were measured and are presented in coefficient form refereed to airplane axes. The results indicate that the change in yawing moment produced by the rudder with the elevator up was the only component of force or moment produced by the elevator and rudder that could not have been balanced in an actual spin by small changes in attitude and angular velocity.

Bamber, M J↗
NASA NTRS1960

Low-Speed Aerodynamic Characteristics of a Model of a Hypersonic Research Airplane at Angles of Attack up to 90 deg for a Range of Reynolds Numbers

Static force tests have been made at low subsonic speeds for a model of a hypersonic research airplane in the Langley high-speed 7- by 10-foot tunnel to determine the aerodynamic forces and moments up to an angle of attack of 90 deg for a range of Reynolds numbers. The Reynolds numbers, based on the mean aerodynamic chord, ranged from 740,000 to 1,900,000, which correspond to dynamic pressures from 15 to 100 lb/sq ft (Mach numbers from 0.10 to 0.27). The model was tested in the clean configuration with various horizontal-tail settings, horizontal tail off, lower rudder off, fuselage alone, and with various size strakes and slats on the nose of the model. Representative results of the present investigation are presented in plotted form, and a tabulation of all the data obtained is presented in a table. Appreciable effects on side force, yawing moment, and pitching moment are indicated by changes in Reynolds number for angles of attack of 40 to 90 deg.

Bowman, James S., Jr.↗
NASA NTRS1976

Effect of Krueger nose flaps on the experimental force and moment characteristics of an oblique wing

Experimental force and moment data are presented for an oblique wing mounted on a body of revolution and equipped with Krueger type nose flaps. The effectiveness of these flaps in making the moment curves more linear by controlling the flow separation on the downstream wing panel at high lift coefficients was determined. The investigation of the effects of the Krueger flaps covered two cases: (1) use of the flaps on the downstream wing panel only and (2) use of the flaps on both wing panels. For part of the tests, the Krueger flaps were mounted on nose flaps that were drooped either 5 deg or 10 deg. The wing was elliptical in planform, had an aspect ratio of 6.0 (based on the unswept span) and was tested at sweep angles of 0, 45 deg, and 50 deg. The Mach-number range covered was from 0.25 to 0.95. It was found that the most effective arrangement of the Krueger flaps for making the pitching-, rolling-, and yawing-moment curves more linear at high lift coefficients was having the Krueger flaps mounted on the nose flaps drooped 5 deg and only on the downstream wing panel.

Hopkins, E. J.↗
DOE OSTI2026

Idealized Offshore Low‐Level Jets for Turbine Structural Impact Considerations

Low-level jets (LLJs) describe conditions in which the wind speed reaches a local maximum with respect to altitude near the surface; they have been observed intermittently in the US mid-Atlantic offshore environment. LLJs pose unique operating conditions for future wind turbines operating in the region, such as negative shear and locally strong veer, but they are not typically considered in existing turbine standards. This work builds upon recent research that explains the formation and evolution of the US mid-Atlantic LLJs through a simple analytical governing equation. We generate several LLJ inflow conditions with varying jet characteristics based on this analytical model and create monotonically sheared (MS) analogues with matched veer to assess the impacts of the LLJ on turbine performance and loading. Using aeroelastic simulations with these inflow conditions on the International Energy Agency 15-MW reference turbine, we find that the LLJ leads to a greater range of tower-top pitching and yawing moments, which could contribute to larger accumulated structural fatigue in components compared to MS inflow. These preliminary results demonstrate a path toward a unified set of test cases for low-level wind maxima that can inform the International Electrotechnical Commission standards related to offshore wind turbine design.

17 WIND ENERGY↗
NASA NTRS1972

Aerodynamic characteristics of some modified conical bodies with low lift-drag ratios at Mach numbers of 2.30 and 4.63

A wind-tunnel investigation was conducted at Mach numbers of 2.30 and 4.63 to determine the static aerodynamic characteristics of three 60 deg half-angle cone models. Configuration 1 was obtained by raking off a symmetrical cone at a base angle of 6.15 deg, and configuration 2 and 3 were obtained by adding flaps to a symmetrical cone. The models were tested at angles of attack from about -5 deg to about 20 deg at roll angles of 0 deg to -180 deg and at a freestream Reynolds number of 1.09 x one million, based on body diameter. The results showed that all three configurations produced finite values of lift-drag ratio useful for lifting planetary entry. All three configurations exhibited increases in yawing moment and side force with roll angle; thus, the capability for lateral trajectory control is provided.

Davenport, E. E.↗
NASA NTRS1972

Effect of configuration modification on the hypersonic aerodynamic characteristics of a blended delta wing-body entry vehicle

The longitudinal, lateral, and directional aerodynamic characteristics of a delta-wing configuration were obtained experimentally at Mach 20 in helium with Reynolds numbers, based on model length, of 1.5 million and 2.9 million and at a Mach number of 6 in air with a Reynolds number, based on model length, of 4.8 million. The angles of attack varied from 0 deg to 55 deg for two sideslip angles. The effects of the addition of dorsal fins, the removal of wing tip fins, an increase in elevon span, and changes in elevon hinge-line sweep angle are discussed. The unmodified vehicle had a maximum lift-drag ratio of 2.1 at Mach 19 and of 2.4 at Mach 6 with about the same lateral and directional stability level at both Mach numbers. As the Mach number increased from 6 to 20, the longitudinal center of pressure moved forward and more positive elevon deflection was therefore required to maintain a given trim angle. The removal of wing tip fins increased the maximum lift-drag ratio and had a negligible effect on longitudinal stability, but caused directional instability that was not corrected by the dorsal fins examined. The shape of the wing and elevon hinge-line sweep had a large influence on the induced yawing moment due to roll control.

Arrington, J. P.↗
NASA NTRS1976

Wind tunnel investigation of a 14 foot vertical axis windmill

A full scale wind tunnel investigation was made to determine the performance characteristics of a 14 ft diameter vertical axis windmill. The parameters measured were wind velocity, shaft torque, shaft rotation rate, along with the drag and yawing moment. A velocity survey of the flow field downstream of the windmill was also made. The results of these tests along with some analytically predicted data are presented in the form of generalized data as a function of tip speed ratio.

Muraca, R. J.↗
NASA NTRS1977

Side forces on forebodies at high angles of attack and Mach numbers from 0.1 to 0.7: two tangent ogives, paraboloid and cone

An experimental investigation was conducted in the Ames 12-Foot Wind Tunnel to determine the subsonic aerodynamic characteristics of four forebodies at high angles of attack. The forebodies tested were a tangent ogive with fineness ratio of 5, a paraboloid with fineness ratio of 3.5, a 20 deg cone, and a tangent ogive with an elliptic cross section. The investigation included the effects of nose bluntness and boundary-layer trips. The tangent-ogive forebody was also tested in the presence of a short afterbody and with the afterbody attached. Static longitudinal and lateral/directional stability data were obtained. The investigation was conducted to investigate the existence of large side forces and yawing moments at high angles of attack and zero sideslip. It was found that all of the forebodies experience steady side forces that start at angles of attack of from 20 deg to 35 deg and exist to as high as 80 deg, depending on forebody shape. The side is as large as 1.6 times the normal force and is generally repeatable with increasing and decreasing angle of attack and, also, from test to test. The side force is very sensitive to the nature of the boundary layer, as indicated by large changes with boundary trips. The maximum side force caries considerably with Reynolds number and tends to decrease with increasing Mach number. The direction of the side force is sensitive to the body geometry near the nose. The angle of attack of onset of side force is not strongly influenced by Reynolds number or Mach number but varies with forebody shape. Maximum normal force often occurs at angles of attack near 60 deg. The effect of the elliptic cross section is to reduce the angle of onset by about 10 deg compared to that of an equivalent circular forebody with the same fineness ratio. The short afterbody reduces the angle of onset by about 5 deg.

Keener, E. R.↗