SIMILITUDE REQUIREMENTS FOR SCALE MODEL DETERMINATION OF SHELL BUCKLING UNDER IMPULSIVE PRESSURES
Similitude requirements in scale model determination of shell buckling under impulsive pressures
BUCKLING↗
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BASIC PRESSURE MEASUREMENTS ON A 0.0667-SCALE MODEL OF THE NORTH AMERICAN X-15 RESEARCH AIRPLANE AT TRANSONIC SPEEDS
Basic pressure measurements on 0.0667-scale model of north american x-15 research aircraft at transonic speeds
X-15A AIRCRAFT↗
Results of a wind tunnel/flight test program to compare afterbody/nozzle pressures on a 1/12 scale model and an F-15 aircraft
In 1975 NASA Dryden Flight Research Facility received the No. 2 prototype F-15 aircraft from the USAF to conduct the F-15 Propulsion/Airframe Interactions Program. About the same time, NASA Langley Research Center acquired a 1/12 scale F-15 propulsion model, whose size made it suitable for detailed afterbody/nozzle static pressure distribution studies. Close coordination between Langley and Dryden assured identical orifice locations and nozzle geometries on the model and aircraft. This paper discusses the sequence of the test programs and how retesting the model after completion of the flight tests greatly increased the ability to match hardware and test conditions. The experience gained over the past decade from involvement in the program should prove valuable to any future programs attempting to match wind tunnel and flight test conditions and hardware.
Pendergraft, O. C., Jr.↗
Pressure distributions on a 0.02-scale Space Shuttle orbiter nose at Mach 21.5 in helium
Pressure distributions on a 0.02-scale model of the Space Shuttle orbiter forward fuselage were obtained in the 22-inch aerodynamic leg of the Langley Hypersonic Helium Tunnel Facility at a nominal free-stream Mach number of 21.5 and a ratio of specific heats of 1.67 for inclusion in the database of the Shuttle entry air data system (SEADS). The data were measured at model angles of attack of 0 deg to 50 deg in 5 deg increments for zero sideslip angle and at model sideslip angles of -5 deg to 5 deg for angles of attack equal to 5, 20, 35, and 40 deg. These data displayed trends similar to those observed in other wind tunnels at Mach 6 and 10 in air. Specifically noted is a shift in the location of the stagnation point at angles of attack above 15 deg; this effect did not, however, occur in flight. By comparison, the data obtained at Mach 6 in the Langley Hypersonic CF4 Tunnel, corresponding to a lower ratio of specific heats in the postshock region than those in helium and air, showed some reduction of the stagnation point shift at the higher angles of attack. The differences between flight and wind tunnel pressure distributions are believed due primarily to high-temperature gas chemistry effects in flight, which include lower effective specific heat ratios but which were not completely duplicated in the wind tunnels.
Ashby, George C., Jr.↗
Contact conductance evaluation for a full scale space erectable radiator pressurized interface
The baseline thermal control configuration for the Space Station Freedom includes a contact heat exchanger to provide efficient heat transfer between the two-phase thermal bus heat collection/delivery system and the radiator panel heat rejection system. The contact heat exchanger provides a dry interface for a modular radiator system with easy on-orbit panel replacement. July 1988 testing of the Space Erectable Radiator System (SERS) at NASA-JSC provided thermal/vacuum data for three full-scale prototype units of a pressurized dry contact heat exchanger design. Derived contact conductance values agreed with predictions and previous element tests and demonstrated high conductance for relatively low pressure levels. A limited amount of data was also obtained below the operating pressure, resulting in contact conductance trends with respect to interface pressure.
Duschatko, R. John↗
In-flight aeroacoustic environments on prospective space shuttle vehicles
The problem of in-flight aerodynamic noise has been studied by reliable estimates of full scale surface-pressure fluctuations from scale model tests in wind tunnels. Scaling relationships have been verified, and many details of the fluctuating pressure characteristics such as spatial correlation and convection velocities are understood. The effects of the wind tunnel environmental turbulence and noise have also been investigated sufficiently so that threshold levels of usable data are known.
Dods, J. B., Jr.↗
Observations on the fourth-power scaling of high-pressure shock waves in solids
This paper is concerned with the bounds and approximations that are found in the original derivation of the fourth-power scaling of shock waves in solids [D. Grady, Appl. Phys. Lett. 38, 825 (1981)]. The analysis is focused on the framework of the derivation and is independent of constitutive assumption, such as visco-plastic behavior. Results include an upper bound for the shock pressure and a restriction on the range of power coefficient for materials having a power-type function for the shock velocity–particle velocity relation. Relaxation of this restriction is proposed based on the idea that the rise time along the Rayleigh line is strain dependent. The idea led to the application of the Onsager relation to strain rate calculation that in turn resulted in a quadratic function for the strain rate–shock pressure relation. The fourth-power relation is obtained by generalizing the Onsager relation through the introduction of a quadratic dissipative potential in analogy to Rayleigh's dissipative potential.
Horie, Yasuyuki (ORCID:0000000245207246)↗
Ares I Scale Model Acoustic Test Instrumentation for Acoustic and Pressure Measurements
Ares I Scale Model Acoustic Test (ASMAT) is a 5% scale model test of the Ares I vehicle, launch pad and support structures conducted at MSFC to verify acoustic and ignition environments and evaluate water suppression systems Test design considerations 5% measurements must be scaled to full scale requiring high frequency measurements Users had different frequencies of interest Acoustics: 200 - 2,000 Hz full scale equals 4,000 - 40,000 Hz model scale Ignition Transient: 0 - 100 Hz full scale equals 0 - 2,000 Hz model scale Environment exposure Weather exposure: heat, humidity, thunderstorms, rain, cold and snow Test environments: Plume impingement heat and pressure, and water deluge impingement Several types of sensors were used to measure the environments Different instrument mounts were used according to the location and exposure to the environment This presentation addresses the observed effects of the selected sensors and mount design on the acoustic and pressure measurements
Vargas, Magda B.↗
Survey of wing and flap lower-surface temperatures and pressures during full-scale ground tests of an externally blown flap system
Full-scale ground tests of an externally blown flap system were made using the wing of an F-111B airplane and a CF700 engine. Pressure and temperature distributions were determined on the undersurface of the wing, vane, and flap for two engine exhaust nozzles (conical and daisy) at several engine power and engine/wing positions. The tests were made with no airflow over the wing. The leading-edge wing sweep angle was fixed at 26 deg, the angle of incidence between the engine and the wing was fixed at 3 deg, and the tests were conducted with the flap retracted, extended and deflected 35 deg, and extended and deflected 60 deg. The integrated local pressures on the undersurface of the flap produced loads approximately three times as great at the 60 deg flap position as at the 35 deg flap position. With both nozzle configurations, more than 90 percent of the integrated pressure loads were contained within plus or minus 20 percent of the flap span centered around the engine exhaust centerline. The maximum temperature recorded on the flaps was 218 C (424 F) for the conical nozzle and 180 C (356 F) for the daisy nozzle.
Hughes, D. L.↗
Thermal Effects of Pressure Surge and Adiabatic Compression in the Priming of Propulsion and Propellant Transfer Systems
Surge pressures and adiabatic compression during priming of evacuated lines in propulsion and propellant transfer systems has been considered. Test data were obtained for desaturated and saturated water for surge pressures and adiabatic compression for various supply pressures and temperatures. Scaling of surge pressures a t different temperatures has been carried out with the aid of Joukowsky model. The thermal effects associated with surge pressure and the attendant adiabatic compression have been investigated with the a id of a simplified analytical model, which represents an extension of the model developed by Lecourt and Steelant. It has been shown that the predictions from the model compare satisfactorily with test data for a water/helium system, and existing test data for different liquid/gas combinations.
Pressure Surge↗
Wing surface pressure data from large-scale wind-tunnel tests of a propeller-driven STOL model
Wing surface pressure data from large scale wind tunnel tests of propeller driven short takeoff aircraft
Page, V. R.↗
Diffusion Time-Scale of Porous Pressure-Sensitive Paint
Pressure-sensitive paint (PSP) is an optical pressure sensor that utilizes the oxygen quenching of luminescence. PSP measurements in unsteady aerodynamic flows require fast time response of the paint. There are two characteristic time-scales that are related to the time response of PSP. One is the luminescent lifetime representing an intrinsic physical limit for the achievable temporal resolution of PSP. Another is the time-scale of oxygen diffusion across the PSP layer. When the time-scale of oxygen diffusion is much larger than the luminescent lifetime, the time response of PSP is controlled by oxygen diffusion. In a thin homogenous polymer layer where diffusion is Fickian, the oxygen concentration 1021 can be described by the diffusion equation in one-dimension.
Liu, Tianshu↗
Joint Acceptance Attenuation Factor of Integrated Pressure with Unsteady Pressure-Sensitive Paint Measurements
The Unsteady Pressure-Sensitive Paint (uPSP) is widely used to measure the surface pressure of scaled models in wind tunnel tests. Compared to the conventional pressure transducers, uPSP has the advantage of high spatial resolution. With multiple high-speed Complementary Metal Oxide Superconductor (CMOS) cameras, the uPSP data collected with the camera pixels are mapped to the surface grid of the scaled model and converted to pressure. The shot noise is the dominant component of the noise in the uPSP measurement. The integrated pressure is usually computed on the grid nodes of a user-defined patch. The effect of shot noise is reduced in the integrated pressure on the patch; however, the measurement of the aerodynamic pressure may also be attenuated by the decorrelation of the flow pressure field being measured. This paper discusses the Joint Acceptance Attenuation Factor (JAAF) of the integrated pressure with uPSP measurements. The JAAF, a function of frequency, is defined as the ratio of the Power Spectral Density (PSD) of the integrated aerodynamic pressure on the patch to the average PSD of the aerodynamic pressure on the grid nodes of the patch. In this paper, the JAAF is investigated for the integrated pressure on rectangular patches, whose edges are defined in the direction along the streamline or across the streamline. Based on the assumption that the surface pressure field can be described by the Corcos model, the closed-form formulas to compute the JAAFs of the integrated pressure on the discrete grid nodes of a rectangular patch and over the continuous area of a rectangular patch are derived respectively. It is shown that the JAAF of the integrated pressure over the continuous area of a rectangular patch is the limit of that on discrete nodes of the rectangular patch when the number of nodes in each row or column goes to infinity. The closed-form formulas of the JAAF derived in this paper, with estimated parameters of the model, are verified with the measurements of the uPSP and the conventional pressure transducer collected in the Space Launch System Ascent Unsteady Aerodynamics Test at NASA Ames Research Center in November 2017. The closed-form formulas of the JAAF of the integrated pressure on the rectangular patches, based on the Corcos model, provide an efficient method to estimate the attenuation of integration by the decorrelation of the flow pressure field and set references for the comparison of the spectrum of the integrated uPSP measurements and the conventional pressure transducer measurements. The work described in this paper is a part of NASA’s development of a new state-of-the-art uPSP capability in production wind tunnels. Funding was provided by the NASA Aerosciences Evaluation and Test Capabilities Portfolio Office.
acoustics↗
Joint Acceptance Attenuation Factor of Integrated Pressure with Unsteady Pressure-Sensitive Paint Measurements
The Unsteady Pressure-Sensitive Paint (uPSP) is widely used to measure the surface pressure of scaled models in wind tunnel tests. Compared to the conventional pressure transducers, uPSP has the advantage of high spatial resolution. With multiple high-speed Complementary Metal Oxide Superconductor (CMOS) cameras, the uPSP data collected with the camera pixels are mapped to the surface grid of the scaled model and converted to pressure. The shot noise is the dominant component of the noise in the uPSP measurement. The integrated pressure is usually computed on the grid nodes of a user-defined patch. The effect of shot noise is reduced in the integrated pressure on the patch; however, the measurement of the aerodynamic pressure may also be attenuated by the decorrelation of the flow pressure field being measured. This paper discusses the Joint Acceptance Attenuation Factor (JAAF) of the integrated pressure with uPSP measurements. The JAAF, a function of frequency, is defined as the ratio of the Power Spectral Density (PSD) of the integrated aerodynamic pressure on the patch to the average PSD of the aerodynamic pressure on the grid nodes of the patch. In this paper, the JAAF is investigated for the integrated pressure on rectangular patches, whose edges are defined in the direction along the streamline or across the streamline. Based on the assumption that the surface pressure field can be described by the Corcos model, the closed-form formulas to compute the JAAFs of the integrated pressure on the discrete grid nodes of a rectangular patch and over the continuous area of a rectangular patch are derived respectively. It is shown that the JAAF of the integrated pressure over the continuous area of a rectangular patch is the limit of that on discrete nodes of the rectangular patch when the number of nodes in each row or column goes to infinity. The closed-form formulas of the JAAF derived in this paper, with estimated parameters of the model, are verified with the measurements of the uPSP and the conventional pressure transducer collected in the Space Launch System Ascent Unsteady Aerodynamics Test at NASA Ames Research Center in November 2017. The closed-form formulas of the JAAF of the integrated pressure on the rectangular patches, based on the Corcos model, provide an efficient method to estimate the attenuation of integration by the decorrelation of the flow pressure field and set references for the comparison of the spectrum of the integrated uPSP measurements and the conventional pressure transducer measurements. The work described in this paper is a part of NASA’s development of a new state-of-the-art uPSP capability in production wind tunnels. Funding was provided by the NASA Aerosciences Evaluation and Test Capabilities Portfolio Office.
acoustics↗
Shuttle model tailcone pressure distribution at low subsonic speeds of a 0.03614-scale model in the NASA/LaRC low-turbulence pressure tunnel (LA81), volume 1
An investigation was conducted in the NASA/LaRC Low-Turbulence Pressure Tunnel on a 0.03614-scale orbiter model of a 089B configuration with a 139B configuration nose forward of F.S. 500. The tailcone was the TC sub 4 design and was instrumented with eighty-nine pressure orifices. Control surfaces were deflected and three wind tunnel mounting techniques were investigated over an angle-of-attack range from -2 deg to a maximum of 18 deg. In order to determine the sensitivity of the tailcone to changes in Reynolds number, most of the test was made at a Mach number of 0.20 over a Reynolds number range of 2.0 to 10 million per foot. A few runs were made at a Mach number of 0.30 at Reynolds numbers of 4.0, 6.0, and 8 million per foot.
Ball, J. W.↗
A graphics package for meteorological data, version 1.5
A plotting package has been developed to simplify the task of plotting meteorological data. The calling sequences and examples of high level yet flexible routines which allow contouring, vectors and shading of cylindrical, polar, orthographic and Mollweide (egg) projections are given. Routines are also included for contouring pressure-latitude and pressure-longitude fields with linear or log scales in pressure (interpolation to fixed grid interval is done automatically). Also included is a fairly general line plotting routine. The present version (1.5) produces plots on WMS laser printers and uses graphics primitives from WOLFPLOT.
Moorthi, Shrinivas↗
Experimental and analytical study of cryogenic propellant boiloff to develop and verify alternate pressurization concepts for Space Shuttle external tank using a scaled down tank
Self pressurization by propellant boiloff is experimentally studied as an alternate pressurization concept for the Space Shuttle external tank (ET). The experimental setup used in the study is an open flow system which is composed of a variable area test tank and a recovery tank. The vacuum jacketed test tank is geometrically similar to the external LOx tank for the Space Shuttle. It is equipped with instrumentation to measure the temperature and pressure histories within the liquid and vapor, and viewports to accommodate visual observations and Laser-Doppler Anemometry measurements of fluid velocities. A set of experiments were conducted using liquid Nitrogen to determine the temperature stratification in the liquid and vapor, and pressure histories of the vapor during sudden and continuous depressurization for various different boundary and initial conditions. The study also includes the development and calibration of a computer model to simulate the experiments. This model is a one-dimensional, multi-node type which assumes the liquid and the vapor to be under non-equilibrium conditions during the depressurization. It has been tested for a limited number of cases. The preliminary results indicate that the accuracy of the simulations is determined by the accuracy of the heat transfer coefficients for the vapor and the liquid at the interface which are taken to be the calibration parameters in the present model.
Akyuzlu, K. M.↗
Experimental evaluation of shear modulus scaling of dynamic strength at extreme pressures
Recent progress in the development of dynamic strength experimental platforms is allowing for unprecedented insight into the assumptions used to construct constitutive models operating in extreme conditions. In this work, we make a quantitative assessment of how tantalum strength scales with its shear modulus to pressures of hundreds of gigapascals through a cross-platform examination of three dynamic strength experiments. Specifically, we make use of Split–Hopkinson pressure bar and Richtmyer–Meshkov instability experiments to assess the low-pressure strain and strain rate dependence. Concurrent examination of magnetically driven ramp-release experiments up to pressures of 350GPa allows us to examine the pressure dependence. In this work, using a modern description of the shear modulus, validated against both ab initio theory and experimental measurements, we then assess how the experimentally measured pressure dependence scales with shear modulus. Furthermore, we find that the common assumption of scaling strength linearly with the shear modulus is too soft at high pressures and offer discussion as to how descriptions of slip mediated plasticity could result in an alternative scaling that is consistent with the data.
71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗