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At least 1,297 records · Page 72
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

On the design of contractions and settling chambers for optimal turbulence manipulations in wind tunnels

The sensitivity of a moderate axisymmetric contraction of turbulence to upstream conditions is parametrically investigated in terms of turbulence intensities and integral length scales. Semi-empirical correlations are derived to characterize the development of turbulence intensities and integral length scales downstream of typical wind-tunnel turbulence manipulators. A new approach is described, permitting wind-tunnel designers to select the characteristic mesh of the turbulence manipulators, a well as the distance from them to the contraction, in order to obtain certain required turbulence characteristics in the test section at the exit of the contraction. The design charts strongly suggest that it may not be possible to achieve small scales and low intensities simultaneously, when using a contraction ratio of nine or larger. The results also confirm that the rapid distortion theory is only valid, with a proper viscous dissipation correction, up to the 'critical point' of a contraction. At that point, the scales of the various components of the velocity are equal and intercomponent transfer from the lateral to the streamwise velocities commences.

Nagib, H. M.↗
NASA NTRS1990

Using transonic small disturbance theory for predicting the aeroelastic stability of a flexible wind-tunnel model

The CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code, developed at the NASA - Langley Research Center, is applied to the Active Flexible Wing (AFW) wind tunnel model for prediction of the model's transonic aeroelastic behavior. Static aeroelastic solutions using CAP-TSD are computed. Dynamic (flutter) analyses are then performed as perturbations about the static aeroelastic deformations of the AFW. The accuracy of the static aeroelastic procedure is investigated by comparing analytical results to those from previous AFW wind tunnel experiments. Dynamic results are presented in the form of root loci at different Mach numbers for a heavy gas and air. The resultant flutter boundaries for both gases are also presented. The effects of viscous damping and angle-of-attack, on the flutter boundary in air, are presented as well.

Silva, Walter A.↗
NASA NTRS1993

Assessment of a flow-through balance for hypersonic wind tunnel models with scramjet exhaust flow simulation

The purpose of this investigation were twofold: first, to determine whether accurate force and moment data could be obtained during hypersonic wind tunnel tests of a model with a scramjet exhaust flow simulation that uses a representative nonwatercooled, flow-through balance; second, to analyze temperature time histories on various parts of the balance to address thermal effects on force and moment data. The tests were conducted in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel at free-stream Reynolds numbers ranging from 0.5 to 7.4 x 10(exp 6)/ft and nominal angles of attack of -3.5 deg, 0 deg, and 5 deg. The simulant exhaust gases were cold air, hot air, and a mixture of 50 percent Argon and 50 percent Freon by volume, which reached stagnation temperatures within the balance of 111, 214, and 283 F, respectively. All force and moment values were unaffected by the balance thermal response from exhaust gas simulation and external aerodynamic heating except for axial-force measurements, which were significantly affected by balance heating. This investigation showed that for this model at the conditions tested, a nonwatercooled, flow-through balance is not suitable for axial-force measurements during scramjet exhaust flow simulation tests at hypersonic speeds. In general, heated exhaust gas may produce unacceptable force and moment uncertainties when used with thermally sensitive balances.

Huebner, Lawrence D.↗
NASA NTRS1973

A Wind-Tunnel Investigation of the Application of the NASA Supercritical Airfoil to a Variable-Wing-Sweep Fighter Airplane

An investigation was conducted in the Langley 8 foot transonic pressure tunnel and the Langley Unitary Plan wind tunnel to evaluate the effectiveness of three variations of the NASA supercritical airfoil as applied to a model of a variable wing sweep fighter airplane. Wing panels incorporating conventional NACA 64A series airfoil with 0.20 and 0.40 camber were used as bases of reference for this evaluation. Static force and moment measurements were obtained for wing leading edge sweep angles of 26, 33, 39, and 72.5 degrees. Fluctuating wing root bending moment data were obtained at subsonic speeds to determine buffet characteristics. Subsonic data were also obtained for determining the effects of wing transition location and spoiler deflection. Limited lateral directional data are included for the conventional 0.20 cambered wing and the supercritical wing.

Ayers, T. G.↗
NASA NTRS1941

Wind-tunnel investigation of effect of interference on lateral-stability characteristics of four NACA 23012 wings, an elliptical and a circular fuselage and vertical fins

Report presents the results of a wind-tunnel investigation of the effect of wing-fuselage interference on lateral-stability characteristics made in the NACA 7 by 10-foot wind tunnel on four fuselages and two fins, representing high-wing, low-wing, and midwing monoplanes. The fuselages are of circular and elliptical cross section. The wings have rounded tips and, in plan form, one is rectangular and the three are tapered 3:1 with various amounts of sweep. The rate of change in the coefficients of rolling moment, yawing moment, and lateral force with angle of yaw is given in a form to show the increment caused by wing-fuselage interference for the model with no fin and the effect of wing-fuselage interference on fin effectiveness. Results for the fuselage-fin combination and the wing tested alone are also given.

House, Rufus O↗
NASA NTRS1976

Investigation of minimum operating temperatures for cryogenic wind tunnels

Total temperatures corresponding to the onset of condensation effects were determined for flow over a 0.137-m NACA 0012-64 airfoil mounted in the Langley 1/3-m transonic wind tunnel. Tests were carried out at a total pressure range from 1.2 to 4.5 atm and at free-stream Mach numbers of 0.75, 0.85, and 0.95. No condensation effects were found to occur until total temperatures were below those associated with free-stream saturation. Significant increases in Reynolds number may apparently be obtained by operation at wind tunnel temperatures below those associated with local saturation over the airfoil but above those where effects first occur. For the 0.85 and 0.95 Mach numbers the increase in Reynolds number was at least 15% over those achieved at local saturation conditions for the same pressure range. Higher pressures may have had some detrimental effect in the 0.95 Mach number test.

Hall, R. M.↗
NASA NTRS1978

Flight effects on noise by the JT8D engine with inverted primary/fan flow as measured in the NASA-Ames 40 by 80 foot wind tunnel

A JT8D-17R engine with inverted primary and fan flows was tested under static conditions as well as in the NASA Ames 40 by 80 Foot Wind Tunnel to determine static and flight noise characteristics, and flow profile of a large scale engine. Test and analysis techniques developed by a previous model and JT8D engine test program were used to determine the in-flight noise. The engine with inverted flow was tested with a conical nozzle and with a plug nozzle, 20 lobe nozzle, and an acoustic shield. Wind tunnel results show that forward velocity causes significant reduction in peak PNL suppression relative to uninverted flow. The loss of EPNL suppression is relatively modest. The in-flight peak PNL suppression of the inverter with conical nozzle was 2.5 PNdb relative to a static value of 5.5 PNdb. The corresponding EPNL suppression was 4.0 EPNdb for flight and 5.0 EPNdb for static operation. The highest in-flight EPNL suppression was 7.5 EPNdb obtained by the inverter with 20 lobe nozzle and acoustic shield. When compared with the JT8D engine with internal mixer, the inverted flow configuration provides more EPNL suppression under both static and flight conditions.

Strout, F. G.↗
NASA NTRS1969

Heat-Transfer Measurements Obtained on the X-15 Airplane Including Correlations With Wind-Tunnel Results

Heat-transfer measurements were obtained on the X-15 airplane from two flights under quasi-steady conditions at a free-stream Mach number of 5.1 and an angle of attack of 2.0° and a free-stream Mach number of 4.98 and an angle of attack of 16.3°. These measurements were made at corresponding free-stream Reynolds numbers of 2.45 × 10 6 and 1.31 × 10 6 per foot (8.04 × 10 6 and 4.30 × 10 6 per meter). Experimental heat-transfer coefficients derived from temperatures obtained from 200 recording thermocouples on the skin of the airplane are tabulated. Correlations with wind-tunnel results show that the wind-tunnel data are in fair-to-good agreement with the flight data obtained on the wing, ventral tail, and vertical tail at low angles of attack but are generally in poor agreement with the high-angle-of-attack data and the low-angle-of-attack fuselage data.

Robert D Quinn↗
NASA NTRS1986

Methods for assessing wall interference in the 2- by 2-foot adaptive-wall wind tunnel

Discussed are two methods for assessing two-dimensional wall interference in the adaptive-wall test section of the NASA Ames 2 x 2-Foot Transonic Wind Tunnel: (1) a method for predicting free-air conditions near the walls of the test section (adaptive-wall methods); and (2) a method for estimating wall-induced velocities near the model (correction methods), both of which methods are based on measurements of either one or two components of flow velocity near the walls of the test section. Each method is demonstrated using simulated wind tunnel data and is compared with other methods of the same type. The two-component adaptive-wall and correction methods were found to be preferable to the corresponding one-component methods because: (1) they are more sensitive to, and give a more complete description of, wall interference; (2) they require measurements at fewer locations; (3) they can be used to establish free-stream conditions; and (4) they are independent of a description of the model and constants of integration.

Schairer, E. T.↗
NASA NTRSDate not supplied

Computational Modeling of Mars Retropropulsion Concepts in the Langley Unitary Plan Wind Tunnel

Future human Mars missions will require powered descent starting at supersonic conditions, something which has never been done before at Mars. Computational powered descent flowfield simulations have been completed at full-scale Mars conditions, but the available ground test data are not appropriate for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles. A test will be conducted in the NASA Langley Unitary Plan Wind Tunnel to begin addressing powered descent aerodynamics risks for large-scale human Mars entry concepts and to identify gaps in computational predictive capabilities. This paper covers pre-test computational flowfield predictions of two different models derived from full-scale reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry. Calculations of the blunt model include variations in nozzle configuration: nozzle location, size, area ratio, and pointing direction. There exist some significant differences between solvers, but some general trends are observed from simulations of the blunt model. First, aerodynamic axial force from the heatshield decreases with increasing thrust due to expanding plume blockage. Second, nozzles that point along the model axis result in lower aerodynamic axial force compared to nozzles that have a radial thrust component. Finally, placing the nozzles further away from the model nose preserves more heatshield axial force with increasing thrust com- pared to nozzles that are closer to the nose. For the slender model, the axial force from the heatshield is similar to the non-blowing axial force regardless of thrust magnitude, due to the nozzle arrangement on the heatshield. Once the test is completed, direct comparisons between the computations and test data will be made to determine computational uncertainties in a wind tunnel environment, to identify gaps in predictive capabilities, and to inform planning for future ground and flight test programs for Mars powered descent vehicles.

Mars↗
NASA NTRS1959

A Hydrogen Peroxide Hot-Jet Simulator for Wind-Tunnel Tests of Turbojet-Exit Models

A turbojet-engine-exhaust simulator which utilizes a hydrogen peroxide gas generator has been developed for powered-model testing in wind tunnels with air exchange. Catalytic decomposition of concentrated hydrogen peroxide provides a convenient and easily controlled method of providing a hot jet with characteristics that correspond closely to the jet of a gas turbine engine. The problems associated with simulation of jet exhausts in a transonic wind tunnel which led to the selection of a liquid monopropellant are discussed. The operation of the jet simulator consisting of a thrust balance, gas generator, exit nozzle, and auxiliary control system is described. Static-test data obtained with convergent nozzles are presented and shown to be in good agreement with ideal calculated values.

Runckel, Jack F.↗
NASA NTRS1959

A Hydrogen Peroxide Hot-Jet Simulator for Wind-Tunnel Tests of Turbojet-Exit Models

A turbojet-engine-exhaust simulator which utilizes a hydrogen peroxide gas generator has been developed for powered-model testing in wind tunnels with air exchange. Catalytic decomposition of concentrated hydrogen peroxide provides a convenient and easily controlled method of providing a hot jet with characteristics that correspond closely to the jet of a gas turbine engine. The problems associated with simulation of jet exhausts in a transonic wind tunnel which led to the selection of a liquid monopropellant are discussed. The operation of the jet simulator consisting of a thrust balance, gas generator, exit nozzle, and auxiliary control system is described. Static-test data obtained with convergent nozzles are presented and shown to be in good agreement with ideal calculated values.

Runckel, Jack F.↗