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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 721 records · Page 40
All sourcesNASA NTRSDOE OSTI
NASA NTRS2022

Structural Shape Sensing from Strain Data Using the Basis Function Method

An aircraft shape must be measured during flight to implement an active trim shape control, such as an active dihedral control for a highly flexible aircraft with a high-aspect-ratio wing, and an active sonic boom control for a supersonic transport aircraft. A basis function method for linear as well as geometrically nonlinear structural shape sensing is proposed in this study. Basis functions can be mode shapes, linear static deformation shapes, or geometrically nonlinear static deformation shapes. The proposed basis function method is validated using a high-aspect-ratio wing, a swept test plate, and the National Aeronautics and Space Administration Low Boom Flight Demonstration mission X-59 Quiet Supersonic Technology (QueSST) aircraft (Lockheed Martin, Bethesda, Maryland). A large structural deformation of a high-aspect-ratio wing is successfully captured using the proposed basis function method with less than 0.3-percentprediction error at the wing-tip section. The basis function method gives excellent deformation prediction, even with stress concentration. Performance of the basis functions is compared using the X-59 stabilator sample case. In most of the sample load cases, static deformations give a better correlation with the target deformation than do mode shapes. Wing shape sensing with sparse strain data is also demonstrated in this study using the X-59 QueSST aircraft. Predicted structural deformations match reasonably with the target deformations even without strain gauges on some structural components. The predicted deformations have a good match with the target deformations.

Chan-gi Pak↗
NASA NTRS1973

Canard-wing lift interference related to maneuvering aircraft at subsonic speeds

An investigation was conducted at Mach numbers of 0.7 and 0.9 to determine the lift interference effect of canard location on wing planforms typical of maneuvering fighter configurations. The canard had an exposed area of 16.0 percent of the wing reference area and was located in the plane of the wing or in a position 18.5 percent of the wing mean geometric chord above the wing plane. In addition, the canard could be located at two longitudinal stations. Two different wing planforms were tested: one with a leading-edge sweep angle of 60 deg and the other with a leading-edge sweep angle of 44 deg. The results indicated that although downwash from the canard reduced the wing lift at angles of attack up to approximately 16 deg, the total lift was substantially greater with the canard on than with the canard off. At angles of attack above 16 deg, the canard delayed the wing stall. Changing canard deflection had essentially no effect on the total lift, since the additional lift generated by the canard deflection was lost on the wing due to an increased downwash at the wing from the canard.

Gloss, B. B.↗
NASA NTRS2010

20 Plus Years of Chimera Grid Development for the Space Shuttle. STS-107, Return to Flight, End of the Program

This slide presentation reviews the progress in grid development for the space shuttle, with particular focus on the development from the los of STS-107 and the return to flight, to the end of the program. Included are views from the current Space Shuttle Launch Vehicle (SSLV) grid system, containing 1.8 million surface points, and 95+ million volume points. Charts showing wind tunnel tests comparisons, and Computational fluid dynamics (CFD) vs 1A613B wing pressures, wind tunnel test comparison with CFD of the proposed ice/frost ramp configuration are shown. The use of pressure sensitive paint and particle imaging velocimetry was used to support debris transport tools, The actual creation of the grids and the use of overset CFD to assess the external tank redesign was also reviewed. It also asks was the use of the overset tool the right choice. The presentation ends with a review of the work to be done still.

Gomez, Reynaldo J., III↗
NASA NTRSDate not supplied

NASA 5.2%-Scale High Lift Common Research Model (CRM-HL) Test in the National Transonic Facility (NTF)

Problem: Predicting CL,max and assessing transition, turbulence models, and Reynolds number effect predictions in CFD requires wind tunnel data for comparison. Objective: Expand the CFD validation database by running multiple models designed using the same reference geometry in various wind tunnels around the world. The wind tunnel data and model geometry will be open source to allow for CFD validation use. Approach: Run the NASA 5.2%-scale High Lift Common Research Model (CRM-HL) semi-span model at the National Transonic Facility (NTF). Test the model at chord Reynolds numbers (Rec) between 1.6 million and 30 million and at varying Q/Es to assess Reynolds number effects and aeroelasticity affects on the model, while collecting force and moment, pressure and wing deformation data. Test the model with ice shapes installed on the leading edges. Results: Ran eight model configurations, including four landing, two landing with ice shapes installed, and two takeoff configurations. Ran in air (120°F) and nitrogen operations (-50°F, -180°F and -250°F) at 7 different chord Reynolds numbers. The data compare well with previous tests of the same model at DLR (The Germany Aerospace Center) and the 14- by 22-Foot Subsonic Tunnel (14x22). Significance: Provided valuable high Reynolds number data on a high lift configuration to the worldwide research community. Provide icing data at flight Reynolds number to the research community.

CRM-HL↗
NASA NTRS2005

Tests Narrowed

The space shuttle program is preparing three thermal protection system and wing leading edge repair tests that the STS-114 return-to-flight astronaut crew would perform during their mission on board Discovery, tentatively scheduled for mid-May.

Covault, Craig↗
NASA NTRS1982

Dynamic structural aeroelastic stability testing of the XV-15 tilt rotor research aircraft

For the past 20 years, a significant effort has been made to understand and predict the structural aeroelastic stability characteristics of the tilt rotor concept. Beginning with the rotor-pylon oscillation of the XV-3 aircraft, the problem was identified and then subjected to a series of theoretical studies, plus model and full-scale wind tunnel tests. From this data base, methods were developed to predict the structural aeroelastic stability characteristics of the XV-15 Tilt Rotor Research Aircraft. The predicted aeroelastic characteristics are examined in light of the major parameters effecting rotor-pylon-wing stability. Flight test techniques used to obtain XV-15 aeroelastic stability are described. Flight test results are summarized and compared to the predicted values. Wind tunnel results are compared to flight test results and correlated with predicted values.

Schroers, L. G.↗
NASA NTRS1989

Integration effects of pylon geometry on a high-wing transport airplane

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the installation effects of a series of pylons that had differing cross-sectional shapes on the pressure distributions and aerodynamic characteristics of a 1/24-scale high wing transport. The tests were conducted at Mach numbers at 0.70 and 0.80 at angles of attack from -3 degrees to 4 degrees with the pylons tested at various toe angles between 5 degrees inboard and 5 degrees outboard. Results of this study indicate that the installed drag was lowest for the pylons with a compression pylon type design which kept the flow under the wing in the pylon/wing junction comparable to the clean wing velocities.

Carlson, John R.↗
NASA NTRS1989

Capabilities of wind tunnels with two-adaptive walls to minimize boundary interference in 3-D model testing

An initial wind tunnel test was made to validate a new wall adaptation method for 3-D models in test sections with two adaptive walls. First part of the adaptation strategy is an on-line assessment of wall interference at the model position. The wall induced blockage was very small at all test conditions. Lift interference occurred at higher angles of attack with the walls set aerodynamically straight. The adaptation of the top and bottom tunnel walls is aimed at achieving a correctable flow condition. The blockage was virtually zero throughout the wing planform after the wall adjustment. The lift curve measured with the walls adapted agreed very well with interference free data for Mach 0.7, regardless of the vertical position of the wing in the test section. The 2-D wall adaptation can significantly improve the correctability of 3-D model data. Nevertheless, residual spanwise variations of wall interference are inevitable.

Rebstock, Rainer↗
NASA NTRS1942

Supplementary Spin Tests of a 1/20-Scale Model of the Bell XP-39E Airplane

As a result of the loss of the first XP-39E airplane due to an uncontrolled spin, and at the request of the War Department received in their letter to the NACA, March 31, 1942, additional tests have been made of the corresponding 1/20-scale spin-tunnel model. Tests were made for the normal full load condition and also for the loading condition corresponding to that of the airplane at the time of the crash, that Is, with guns, ammunition, and half of the fuel removed from the wings. The model tests were made with the tail arrangement corresponding to that of the airplane at the time of the crash, as well as with numerous modified tail and contour arrangements In an attempt to evolve satisfactory recovery characteristics. The tests were started on April 2, 1942, and completed in approximately two weeks, with the development of several alternate tail arrangements giving reasonably satisfactory recoveries for normal spinning control configuration. The tests were witnessed by Mr. Melvin Shorr, of the Materiel Center at Wright Field, and Mr. Paul C. Emmons, of the Bell Aircraft Corporation, to whom records of the test data were furnished. Colonel H. P. Swofford, of the Materiel Center, Wright Field, also witnessed part of the tests. A preliminary report of the results of the tests has been forwarded to the Materiel Command at Wright Field. A brief check of the parachute installation for the airplane was also made.

A I Neihouse↗
NASA NTRS1975

Flow visualization of vortices locked by spanwise blowing over wings featuring a unique leading and trailing-edge flap system

Flow visualization studies were conducted to qualitatively determine the effects of active generation and augmentation of vortex flow over wings by blowing a discrete jet in a spanwise direction in the channel formed by extension of upper surface leading- and trailing-edge flaps. Spanwise blowing from a reflection plane over a rectangular wing was found to generate and lock a dual corotating vortex system within the channel and, at sufficient blowing rates, cause the separated flow off the upper end of the leading-edge flap to reattach to the trailing-edge flap. Test parameters included wing angle of attack, jet momentum coefficient, leading- and trailing-edge flap deflection angle, and jet location above the wing surface. Effects due to removal of the leading- and trailing-edge flap were also investigated.

Erickson, G. E.↗
NASA NTRS1979

Effects of wing leading-edge deflection on low-speed aerodynamic characteristics of a low-aspect-ratio highly swept arrow-wing configuration

Static force tests were conducted in the Langley V/STOL tunnel at a Reynolds number (based on the mean aerodynamic chord) of about 2.0 x 10 to the 6th power for an angle-of-attack range from about - 10 deg to 17 deg and angles of sideslip of 0 and + or - 5 deg. Limited flow visualization studies were also conducted in order to provide a qualitative assessment of leading-edge upwash characteristics.

Coe, P. L., Jr.↗
NASA NTRS1992

Flight test results from a supercritical mission adaptive wing with smooth variable camber

The mission adaptive wing (MAW) consisted of leading- and trailing-edge variable-camber surfaces that could be deflected in flight to provide a near-ideal wing camber shape for any flight condition. These surfaces featured smooth, flexible upper surfaces and fully enclosed lower surfaces, distinguishing them from conventional flaps that have discontinuous surfaces and exposed or semiexposed mechanisms. Camber shape was controlled by either a manual or automatic flight control system. The wing and aircraft were extensively instrumented to evaluate the local flow characteristics and the total aircraft performance. This paper discusses the interrelationships between the wing pressure, buffet, boundary-layer and flight deflection measurement system analyses and describes the flight maneuvers used to obtain the data. The results are for a wing sweep of 26 deg, a Mach number of 0.85, leading and trailing-edge cambers (delta(sub LE/TE)) of 0/2 and 5/10, and angles of attack from 3.0 deg to 14.0 deg. For the well-behaved flow of the delta(sub LE/TE) = 0/2 camber, a typical cruise camber shape, the local and global data are in good agreement with respect to the flow properties of the wing. For the delta(sub LE/TE) = 5/10 camber, a maneuvering camber shape, the local and global data have similar trends and conclusions, but not the clear-cut agreement observed for cruise camber.

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