Jet-boundary corrections to the downwash behind powered models in rectangular wind tunnels with numerical values for 7- by 10-foot closed wind tunnels
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Dynamic response of a hammerhead launch vehicle wind-tunnel model
A wind-tunnel test of a 1/10th-scale Atlas-Centaur I large payload fairing launch vehicle model was conducted in the NASA Langley Transonic Dynamics Tunnel. The wind tunnel model was an aeroelastically-scaled version of the flight vehicle and was capable of simulating either of the first two bending vibration modes of the full-scale vehicle by a partial mode technique. The primary purpose of the test was to gather data concerning buffet response which could be used to clear the vehicle for flight. Additionally, angle-of-attack studies were conducted and several payload fairing configurations were tested to assess the buffet response and dynamic stability of off-design flight conditions and geometric parameters. No dynamic instabilities were found for any of the configurations tested. The buffet response data for the nominal flight configuration indicate that the unsteady buffet loads represent 5 to 10 percent of the total design load; therefore, the buffet loads are not a large factor affecting the overall vehicle design. Payload fairing length-to-diameter ratio variations were found to have small effects on the buffet response of the model, except in the case of the smallest length-to-diameter models for the second bending mode simulation. The effects of angle of attack on buffet response were found to be small. The model was more sensitive to Mach number changes than to angle of attack. The buffet response results from this wind tunnel test were influenced by the tunnel facility vibration levels. An attempt was made to experimentally reduce the effect of the facility mechanical vibration for the nominal flight configuration by testing with vertical rods used to stiffen the sting support. The first flight of the Atlas-Centaur I vehicle successfully occurred on July 25, 1990, and a comparison of flight measurements with wind tunnel data is presented. The flight data was found to be well within the 3 sigma level of the wind tunnel data.
Cole, Stanley R.↗
Rocket Plume Scaling for Orion Wind Tunnel Testing
A wind tunnel test program was undertaken to assess the jet interaction effects caused by the various solid rocket motors used on the Orion Launch Abort Vehicle (LAV). These interactions of the external flowfield and the various rocket plumes can cause localized aerodynamic disturbances yielding significant and highly non-linear control amplifications and attenuations. This paper discusses the scaling methodologies used to model the flight plumes in the wind tunnel using cold air as the simulant gas. Comparisons of predicted flight, predicted wind tunnel, and measured wind tunnel forces-and-moments and plume flowfields are made to assess the effectiveness of the selected scaling methodologies.
Brauckmann, Gregory J.↗
Potential Bedforms on Venus: Wind Tunnel Simulations
Wind tunnel studies involving mixtures of particle sizes (bimodal and multimodal distributions) indicate the microdunes and related ridge- and wave-like structures can occur over a wider range of conditions. The results indicate that the bedforms develop as a consequence of the effective sorting capability of the high-density Venusian atmosphere. Four mixtures of different particle-size distributions were investigated. The results of these experiments suggest that: (1) small-scale bedforms on Venus may be more common than previously anticipated from simulations involving only unimodal sands; (2) small particles are rapidly sorted on Venus if winds slightly above threshold are available; (3) coarse particles are transported in traction within the bedforms as well as rolled by saltating fines, suggesting that the capacity of wind on Venus to transport material is greater than anticipated from previous flux studies and (4) microdunes and related ridges and waves with coarse and fine layers can be produced during sorting.
Bougan, S. J.↗
A one-component, magnetic support-and-balance system for wind tunnel models.
Wind tunnel model one component magnetic support and balance system for sphere drag investigation at subsonic Mach numbers
Sivier, K. R.↗
Technique for the integral casting of pressure instrumentation in wind-tunnel models
Wind tunnel models are cast around core consisting of array of tubing. Principal advantage of technique is that greater number of pressure orifices are easily installed, without compromising aerodynamic shape of model. Technique reduces construction cost by about 50 percent.
Jackson, C. M., Jr.↗
Investigation of space shuttle launch vehicle external tank nose configuration effects (model 67-OTS) in the Rockwell International 7 by 7 foot trisonic wind tunnel (IA69)
Wind tunnel aerodynamic investigations were conducted on an 0.015-scale representation of the space shuttle launch configuration. The primary test objectives were to investigate shock wave formation and record the aerodynamic stability and control effects generated by a new external tank nose configuration (MCR 467) at a Mach number of 1.2. Schlieren photographs were taken at angles of attack of -4 deg, 0 deg, and 4 deg, beta = 0 deg with force and pressure data recorded over the alpha range of -4 deg equal to or less than alpha equal to or less than 4 deg at beta = + or - 4 deg. The launch configuration model, consisting of the VL70-00014OA/B Orbiter, the VL78-000041B ET, and the VL77-000036A SRBs, was sting mounted on a 2.5-inch Task type internal balance entering through the ET base region. Wing, body, and base pressure lines for all orifices were routed internally through the model to the sting support system. Parametric variation consisted only of altering the ET nose configuration.
Mennell, R.↗
Results of an investigation of Reynolds effects on integrated vehicle elevon hinge moments and wing panel loads obtained with 0.010 scale model 72 OTS in the Rockwell trisonic wind tunnel (IA141)
Wind tunnel investigations were conducted on an 0.010-scale representation of the VL70-000140C Integrated Space Shuttle Launch Vehicle. The primary test objective was to obtain Reynolds number effects on orbiter elevon hinge moments and wing bending/torsional moments. Launch vehicle aerodynamic force data were also recorded. The elevon hinge moments, wing bending/torsional moments, and vehicle force data were recorded over an angle of attack range of -6 deg to +6 deg, an angle of sideslip range of -6 deg to +6 deg, at Mach numbers of 0.6, 0.975, 1.05 and 1.25. The Reynolds number was varied from a minimum of 4.5 million/foot to a maximum of 11.5 million/foot. The complete integrated configuration was tested with the orbiter elevons set at 0 deg and deflected to 9 deg on the outboard elevon and 10 deg on the inboard elevon. Testing was conducted in the TWT 19.7% porous transonic test section with the model sting mounted through the orbiter base. All aerodynamic force data were obtained from internal strain gage balance located in the orbiter.
Mennell, R. C.↗
Laser velocimeter supersonic and transonic wind tunnel studies
Wind tunnel investigations of a shock-induced separated flow and the flow about a two-dimensional airfoil section at transonic speeds using laser velocimetry techniques are described. Laser velocimeter data are presented and compared with conventional pressure type measurements (pitot and static pressure probe measurements for the shock-induced separation, and surface pressure measurements for the two-dimensional airfoil section). In addition to these data comparisons, a simple optical arrangement for the incorporation of a Bragg cell in a laser velocimeter system is described. The effects of particle concentration gradients within a turbulent boundary layer on laser velocimeter measurements are also demonstrated.
Johnson, D. A.↗
Threshold windspeeds for sand on Mars - Wind tunnel simulations
Wind friction threshold speeds for particle movement were determined in a wind tunnel operating at martian surface pressure with a 95 percent CO2 and 5 percent air atmosphere. The relationship between friction speed and free-stream velocity is extended to the critical case for Mars of momentum thickness Reynolds numbers between 425 and 2000. It is determined that the dynamic pressure required to initiate saltation is nearly constant for pressures between 1 bar and 4 mb for atmospheres of both air and CO2.
Greeley, R.↗
A program to evaluate a control system based on feedback of aerodynamic pressure differentials. Part 2: Data report for phase 1 wind tunnel test
Wind tunnel test data are presented. The purpose of the test was to determine the sensor location for sensitivity to flap deflection and to determine the range required of the sensor. The calculations followed to correct the raw pressure measurements, the corrected measurements, which represent the static pressures along the airfoil surface, and the static pressures tabulated and in coefficient form are described. However, the chordwise locations of the upper and lower surface taps do not coincide. An interpolation routine is performed to find the lower surface pressure coefficients at the upper surface tap locations. The difference, C sub P(lower) - C sub P(upper), is calculated and listed and the data are reorganized with respect to each tap location. A numerical regression performed to quantify the slopes of the graphs allows more accurate comparison of sensitivity and linearity.
Hrabak, R. R.↗
Wind tunnel tests of space shuttle solid rocket booster insulation material in the aerothermal tunnel c
Wind tunnel tests of the space shuttle Solid Rocket Booster Insulation were conducted in the von Karman Gas Dynamics Facility Tunnel C. For these tests, Tunnel C was run at Mach 4 with a total temperature of 1100-1440 and a total pressure of 100 psia. Cold wall heating rates were changed by varying the test article support wedge angle. Selected results are presented to illustrate the test techniques and typical data obtained.
Hartman, A. S.↗
Problems associated with operations and measurement in cryogenic wind tunnels
Cryogenic wind tunnel T'3 under continuous blower operation has been the object of improvements and the installation of auxiliary equipment, dealing in particular with the enlargement of the liquid nitrogen injection reservoir and the hook-up to a fast data acquisition system. Following a brief description of the installation and its functioning, we present the main experimental techniques and the instrumentation used in the cryogenic environment.
Blanchard, A.↗
Cryogenic wind tunnels
Cryogenic wind tunnels and their use for high Reynolds number testing are reviewed. Emphasis is given to operational and aerodynamic testing experience in the NASA Langley 0.3-m Transonic Cryogenic Tunnel. Areas covered include development of test techniques and aerodynamic testing in cryogenic tunnels. The use of advanced testing techniques to increase the value of cryogenic tunnels is recommended. These include adaptive wall test sections, using solid but flexible top and bottom walls, and magnetic suspension and balance systems.
Kilgore, Robert A.↗
Development of pneumatic test techniques for subsonic high-lift and in-ground-effect wind tunnel investigations
Wind tunnel evaluations of two-dimensional high-lift airfoils and of vehicles operating in ground effect near the tunnel floor require special test facilities and procedures. These are needed to avoid errors caused by proximity to the walls and interference from the wall boundary layers. Pneumatic test techniques and facilities were developed for GTRI aerodynamic research tunnels and calibrated to verify that these wall effects had been removed. The modified facilities were then employed to evaluate the aerodynamic characteristics of blown very-high-lift airfoils and of racing hydroplanes operating in ground effect at various levels above the floor. The pneumatic facilities, techniques and calibrations are discussed and typical aerodynamic data recorded both with and without the test-section blowing systems are presented.
Englar, Robert J.↗
Investigation of the Effect of Tip Tanks on the Wing Loading of a Republic F-84 Airplane in the Ames 40- by 80-foot Wind Tunnel
Wind-tunnel tests at low Mach number of a Republic F-84C airplane were conducted to determine by pressure-distribution measurements the air loads on wing-tip tanks and the change in wing load distribution due to the presence of tip tanks. Measurements of the aeroelastic twist of the wing were also obtained. Results are presented in the form of loading coefficient, center-of- pressure location, pitching-moment coefficient, aerodynamic-center location, and aeroelastic twist. The investigation revealed that the redistributions in loading brought about by either the tip tanks or elastic deformation of the wing were relatively small when compared with the chnnges in loading normally associated with the deflection of an aileron.
Hunton, Lynn W.↗
DDS-Suite - A Dynamic Data Acquisition, Processing, and Analysis System for Wind Tunnel Testing
Wind Tunnels have optimized their steady-state data systems for acquisition and analysis and even implemented large dynamic-data acquisition systems, however development of near real-time processing and analysis tools for dynamic-data have lagged. DDS-Suite is a set of tools used to acquire, process, and analyze large amounts of dynamic data. Each phase of the testing process: acquisition, processing, and analysis are handled by separate components so that bottlenecks in one phase of the process do not affect the other, leading to a robust system. DDS-Suite is capable of acquiring 672 channels of dynamic data at rate of 275 MB / s. More than 300 channels of the system use 24-bit analog-to-digital cards and are capable of producing data with less than 0.01 of phase difference at 1 kHz. System architecture, design philosophy, and examples of use during NASA Constellation and Fundamental Aerodynamic tests are discussed.
Burnside, Jathan J.↗
Boundary Condition Study for the Juncture Flow Experiment in the NASA Langley 14x22-Foot Subsonic Wind Tunnel
Because future wind tunnel tests associated with the NASA Juncture Flow project are being designed for the purpose of CFD validation, considerable effort is going into the characterization of the wind tunnel boundary conditions, particularly at inflow. This is important not only because wind tunnel flowfield nonuniformities can play a role in integrated testing uncertainties, but also because the better the boundary conditions are known, the better CFD can accurately represent the experiment. This paper describes recent investigative wind tunnel tests involving two methods to measure and characterize the oncoming flow in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The features of each method, as well as some of their pros and cons, are highlighted. Boundary conditions and modeling tactics currently used by CFD for empty-tunnel simulations are also described, and some results using three different CFD codes are shown. Preliminary CFD parametric studies associated with the Juncture Flow model are summarized, to determine sensitivities of the flow near the wing-body juncture region of the model to a variety of modeling decisions.
Rumsey, C. L.↗