Showing papers on "Freestream published in 1995"

Mixing and Penetration Studies of Sonic Jets in a Mach 2 Freestream

Abstract: A comparison of the penetration and mixing characteristics of three transverse/oblique injector configurations is presented. The three geometries studied include circular transverse, circular oblique, and elliptical transverse injectors, and the crossflow is at Mach 2. Planar Mie scattering images of three near-field flow planes produced substantial information about the flowfield created by each injector. In addition to global flowfield characteristics, the Mie scattering images provided transverse and lateral penetrations for each injector. Instantaneous and time-averaged information concerning the structural organization of the flowfields was obtained. Results demonstrate increasing jet penetration in the transverse direction with increasing jet-to-freestream momentum flux ratio. Penetration of the oblique jet is appreciably less in the near-field compared to the two transverse jets due to the reduced component of momentum in the transverse direction. The transverse elliptic jet appears to spread more quickly in the lateral direction than the other two jets, suggesting that some type of axis-switchin g phenomenon occurs. Large-scale structures at the interface between the jet and freestream fluids are shown for the two transversely oriented jets, while small-scale eddies are prominent in the oblique jet flowfield. Near-field mixing appears dominated by these eddies and the counter-rotating structures that develop in the streamwise direction.

Wing flutter boundary prediction using unsteady Euler aerodynamic method

Abstract: Modifications to an existing three-dimensional, implicit, upwind Euler/Navier-Stokes code (CFL3D Version 2.1) for the aeroelastic analysis of wings are described. These modifications, which were previously added to CFL3D Version 1.0, include the incorporation of a deforming mesh algorithm and the addition of the structural equations of motion for their simultaneous time-integration with the government flow equations. The paper gives a brief description of these modifications and presents unsteady calculations which check the modifications to the code. Euler flutter results for an isolated 45 degree swept-back wing are compared with experimental data for seven freestream Mach numbers which define the flutter boundary over a range of Mach number from 0.499 to 1.14. These comparisons show good agreement in flutter characteristics for freestream Mach numbers below unity. For freestream Mach numbers above unity, the computed aeroelastic results predict a premature rise in the flutter boundary as compared with the experimental boundary. Steady and unsteady contours of surface Mach number and pressure are included to illustrate the basic flow characteristics of the time-marching flutter calculations and to aid in identifying possible causes for the premature rise in the computational flutter boundary.

Numerical Investigation of Supersonic Flows Around a Spiked Blunt Body

Abstract: In supersonic flow, a spike attached to the nose reduces the drag of a blunt body. In this paper, supersonic flows around a spiked blunt body are numerically simulated to examine the effects of the spike length, Mach number, and angle of attack. Three-dimensional thin-layer compressible Navier-Stokes equations are solved using a highresolution upwind scheme with LU-ADI time-integration algorithm. The computed results show that the drag of the spiked blunt body is significantly influenced by the spike length, Mach number, and angle of attack. Scales of the separated region are not significantly influenced by the freestream Mach number. For the spiked blunt body at angle of attack, the flowfield becomes complex with spiral flows. The computed results are in reasonable agreement with experimental data.

Prediction of supersonic inlet unstart caused by freestream disturbances

Abstract: The objective of this article is to report progress toward the development of an Euler analysis procedure for predicting the unstart tolerance of supersonic inlets. As an aid to understanding boundary condition issues, a one-dimensional, linear-analysis procedure was developed and used to analyze inlet unstart behavior. Using these results as a guide, an Euler analysis procedure was extended through the addition of a new bleed boundary condition, a new compressor face boundary condition, and an engine demand model for the simulation of unsteady inlet flows caused by freestream flow disturbances. Five unstart conditions were identified with the Euler analysis of the axisymmetric inlet for both 20- and 90-deg throat bleed configurations. Results show that both increases and decreases in temperature or velocity will unstart the inlet, whereas only pressure decreases will unstart the inlet. It was also found that 90-deg throat bleed improves the unstart tolerance relative to 20-deg throat bleed for freestream pressure decreases, temperature increases, and changes in velocity.

Three-dimensional boundary-layer transition on a swept wing at Mach 3.5

Abstract: Transition on a swept-wing leading-edge model at Mach 3.5 is investigated. Surface pressure and temperature measurements are obtained in the NASA Langley Research Center Supersonic Low-Disturbance Tunnel. For one case, temperature-sensitive paint and a sublimating chemical are used to visualize surface flow features such as transition location. The experimental data are compared with 1) mean-flow results computed as solutions to the thin-layer Navier-Stokes equations and 2) N-factors obtained using the envelope e N method. The experimental and computational results compare favorably in most cases. In particular, N ≃ 13 correlates best with the observed transition location over a range of freestream unit Reynolds numbers and angles of attack. Computed traveling crossflow disturbances with frequencies of 40-60 kHz have the largest N factors, and the surface flow visualizations reveal smooth transition fronts with only faint evidence of stationary crossflow vortices. These results suggest that transition is probably dominated by traveling, rather than stationary, crossflow disturbances for the present model.

Control of shock-wave/boundary-layer interactions by bleed

Abstract: This numerical study investigates the effectiveness of bleed in controlling shock-wave/boundary-layer interactions on a flat plate with a focus on understanding how bleed-hole angle, presence of upstream and downstream bleed holes, and pressure ratio across bleed holes affect structure of barrier shock, surface pressure distribution, and bleed rate (in terms of flow coefficient). The bleed-hole angles investigated are 30 deg slanted and 90 deg normal, which give rise to two different types of barrier shocks. The influence of upstream and downstream bleed holes were investigated by studying the bleed process through an isolated hole and through three holes arranged in tandem along the streamwise direction. The plenum/freestream pressure ratios investigated range from 0.3 to 1.7, which produced choked and unchoked flows in the bleed holes. This study is based on the ensemble-averaged, full compressible Navier-Stokes equations closed by the Baldwin-Lomax model with solutions obtained by an implicit finite volume method on an overlapping Chimera grid.

Flow developments above 50-deg sweep delta wings with different leading-edge profiles

Abstract: The characteristics of flow developments above 50-deg sweep delta wings with different leading-edge profiles are shown by flow visualizations and velocity measurements. The Reynolds number based on freestream velocity and root chord is about 7 x 103. The leading-edge profiles studied include the shapes of square, round, windward surface beveling, leeward surface beveling, and wedge. Based on the velocity data obtained along the leading edges of the delta wings it is noted that the flow angles associated with the separated shear layers vary with the leading-edge profiles studied. This finding infers that varying the leading-edge profile has an impact on the initial development of the separated shear layer, consequently, the formation of leading-edge vortex. Furthermore, it is shown that the leading edge of windward beveling causes the largest leading-edge flow angle and produces the most organized leading-edge vortex.

Structure of Supersonic Turbulent Boundary Layer After Expansion Regions

Abstract: The effects of four expansion regions [centered and gradual (R/6o ~ 50) expansions of both 7 and 14 deg] on a fully developed Mach 3 turbulent boundary layer were investigated. Instantaneous visualizations were made possible by the presence of scalar water condensation in the freestream and its absence in the higher temperature boundary layer. The elongated longitudinal structures previously found in the flat plate boundary layer are present downstream of the expansions. Large-scale structures increase in scale across the expansions. Structure angles also initially increase but are found to return to the flat plate value 10#o downstream of the 7-deg centered expansion. The rapid quenching of small-scale turbulence by the expansions results in a more intermittent boundary layer visually dominated by large-scale structures. Convection velocities derived from double-pulse correlations are reasonable in the flat plate and 7-deg centered expansion boundary layers. Excess condensation downstream of the 14-deg expansions (probably CC>2) made the 14-deg expansion results more difficult to interpret. Nomenclature n = normal distance above the surface R = radius of curvature for the gradual expansions, correlation coefficient Ree = Reynolds number based on boundary layer momentum thickness s = streamwise distance along the surface measured from the start of the convex curvature U = mean velocity vector U = mean streamwise velocity UT = friction velocity V = mean normal velocity jc = horizontal distance measured from (s, n) = (0, 0) y = vertical distance measured from (s, n) = (0, 0) Ap = pressure difference across the expansion region SQ = boundary layer thickness at s = 0 mm <$vis = boundary layer thickness defined by 99% of the freestream intensity SRMS = normal distance above the boundary where the peak in the rms profile occurs 9 = boundary layer momentum thickness v = kinematic viscosity TO = surface shear stress ahead of the expansion region

Flow Physics of Critical States for Rolling Delta Wings

Abstract: A delta wing rolling at high flow inclination has two well-documented critical states One occurs when the breakdown of the leeside leading-edge vortex passes over the trailing edge, and the other when the breakdown of the windward vortex reaches the apex The flow physics associated with those critical states are described for a sharp-edged 65-deg delta wing rolling around an axis inclined 30 deg to the freestream It is shown how the measured, extremely nonlinear, unsteady aerodynamics result from the roll-rate-induced camber effect in conjunction with convective flow time lag effects

Effects of chemistry on blunt-body wake structure

Abstract: Results of a numerical study are presented for hypersonic low-density flow about a 70-deg blunt cone using direct simulation Monte Carlo (DSMC) and Navier-Stokes calculations. Particular emphasis is given to the effects of chemistry on the near-wake structure and on the surface quantities and the comparison of the DSMC results with the Navier-Stokes calculations. The flow conditions simulated are those experienced by a space vehicle at an altitude of 85 km and a velocity of 7 km/s during Earth entry. A steady vortex forms in the near wake for these freestream conditions for both chemically reactive and nonreactive air gas models. The size (axial length) of the vortex for the reactive air calculations is 25% larger than that of the nonreactive air calculations. The forebody surface quantities are less sensitive to the chemistry than the base surface quantities. The presence of the afterbody has no effect on the forebody flow structure or the surface quantities. The comparisons of DSMC and Navier-Stokes calculations show good agreement for the wake structure and the forebody surface quantities.

Axisymmetric oscillations of an opposing jet from a hemispherical nose

Abstract: The flowfield around a hemispherical nose with a sonic opposing jet in a freestream of Mach 2.5 is computed by a time-accurate algorithm using the axisymmetric Navier-Stokes equations. Five total pressure ratios of a jet to a freestream, 0.816 ∼ 1.633, are examined. The results obtained strongly depend on the total pressure ratio as the previous experimental results did. In some cases, the computed flowfield oscillates intensely in which the drastic change of the jet structure and the surrounding pressure takes place. The flowfield changes from the stable flow to the unstable flow for the same total pressure ratio as in the experiment The mechanism of the oscillations and the condition of transition between both types of flow are interpreted physically on the basis of unsteady numerical solutions here.

On the Interaction Between the Shock Wave Attached to a Wedge and Freestream Disturbances

Abstract: We present a study of the interaction of small amplitude, unsteady, freestream disturbances with a shock wave induced by a wedge in supersonic flow. These disturbances may be acoustic waves, vorticity waves, or entropy waves (or indeed a combination of all three). Their interactions then generate behind the shock disturbances of all three classes, an aspect that is investigated in some detail. Also, the possibility of enhanced mixing owing to additional vorticity produced by the shock-body coupling is investigated. It is shown that disturbances behind the shock may either decay downstream, or alternatively experience sustained oscillations. The precise regimes under which either behaviour is found are stated.

Laminar Boundary-Layer, Sound Receptivity and Control

Abstract: The receptivity of a laminar boundary layer on a flat plate is experimentally studied in the presence of 2-D roughness and freestream sound. The work is carried out in the ASU Unsteady Wind Tunnel. It is shown that the T-S wave amplitude can be less than that of the smooth-surface case downstream from the roughness due to a superposition of unstable waves. Thus, receptivity studies of localized sources must consider constructive-destructive interference.

Aerothermodynamic Measurement and Prediction for Modified Orbiter at Mach 6 and 10

Abstract: Detailed heat-transfer rate distributions measured laterally over the windward surface of an orbiter-like configuration using thin-film resistance heat-transfer gauges and globally using the newly developed relative intensity, two-color thermographic phosphor technique are presented for Mach 6 and 10 in air. The angle of attack was varied from 0 to 40 deg, and the freestream Reynolds number based on the model length was varied from 4 x 10(exp 5) to 6 x 10(exp 6) at Mach 6, corresponding to laminar, transitional, and turbulent boundary layers; the Reynolds number at Mach 10 was 4 x 10(exp 5), corresponding to laminar flow. The primary objective of the present study was to provide detailed benchmark heat-transfer data for the calibration of computational fluid-dynamics codes. Predictions from a Navier-Stokes solver referred to as the Langley aerothermodynamic upwind relaxation algorithm and an approximate boundary-layer solving method known as the axisymmetric analog three-dimensional boundary layer code are compared with measurement. In general, predicted laminar heat-transfer rates are in good agreement with measurements.

Skin-friction measurements in three-dimensional, supersonic shock-wave/boundary-layer interaction

Abstract: The experimental documentation of a three-dimensional shock-wave/boundary-layer interaction in a nominal Mach 3 flow is presented. The model consisted of a sting-supported cylinder, aligned with the freestream flow, and a 20-deg half-angle conical flare offset 1.27 cm from the cylinder centerline. Surface oil flow, laser light sheet illumination, and spark schlieren photography were used to document the flow topology. Extensive surface-pressure and skin-friction measurements were made throughout the interaction region. A laser interterometric skin-friction instrument was employed to acquire the skin-friction data. Resolved skin-friction measurements of C fx and C fz were made within the highly swept three-dimensional separated regions. The skin-friction data will be of particular value for turbulence modeling and computational fluid dynamics validation

Direct simulation Monte Carlo calculation of a jet interaction experiment

Abstract: This study deals with the development of a methodology for numerically simulating the interaction of a reaction control system (RCS) jet with a low-density external flow. A European Space Agency experiment was chosen as a test case, since it provided experimental data that could validate some of the numerical results. The initial approach was to focus on several subproblems having direct relevance to the full interaction problem. This enabled different numerical methods to be investigated separately and validated for each part of the interaction problem. In this manner, the best methodology for solving the full interaction problem was developed. The subproblems considered in this study included typical RCS nozzle and plume flows, a flat plate at zero incidence, and the flow past the experimental test model without the control jet firing. Once these calculations were completed, a simulation was performed of the test model with the control jet operating at the experimental density. The results from this final simulation were compared with experimental measurements. Nomenclature = reference diameter, m = Knudsen number at nozzle throat = Knudsen number at nozzle lip = Mach number = number density, molecules/m3 = freestream number density, molecules/m3 = pressure, Pa = Reynolds number T^ef = reference temperature, K 7"waii = temperature of nozzle wall T^ = freestream temperature, K X, Y, Z = Cartesian body coordinates p , - density, kg/m3 Poo = freestream density, kg/m3 Kn\ip M

Receptivity mechanisms for Görtler vortex modes

Abstract: The receptivity problem for Gortler vortices induced by wall roughness or freestream disturbances is reviewed. To date, receptivity studies for this problem have been exclusively linear in character and here we show how the roughness and freestream disturbance mechanisms can each play dominant or inconsequential roles as possible routes to transition. The importance of each process is dependent on the exact situation at hand. For example, distributed wall-roughness elements tend to be more important in the generation of O(1) wave-number vortices than are isolated roughness patches whilst variations in the freestream velocity can easily provoke high wave-number disturbances on which roughness distributions typically have negligible effect. It has only been in recent times that the influence a spanwise component of the underlying basic boundary-layer flow may have on the Gortler mechanism has come to be appreciated. In some new computations we show that the imposition of such a spanwise component can lead to an increase in the coupling coefficient (a measure of the efficiency of a generating process) for modes provoked by wall roughness. However, such crossflow tends to reduce the global amplification rate of the most unstable mode so has the overall effect of restricting vortex growth downstream of any roughness element. We also demonstrate how the nonparallelism of Gortler vortices implies that conclusions concerning vortex receptivity properties can only be drawn after taking full account of upstream conditions and the precise form of the generating mechanism but it appears that for a large class of flows distributed wall forcing is more important in the provocation of modes than are either isolated roughness or freestream disturbances.

Hypersonic jet control effectiveness

Abstract: The present study aims to identify some of the parameters which determine the upstream extent and the lateral spreading of the separation front around an under-expanded transverse jet on a slender blunted cone. The tests were conducted in the Cranfield hypersonic facility at $M_\infty = 8.2$ , $Re_\infty /{\rm cm} = 4.5$ to $9.0 \times 10^4$ and at $M_\infty = 12.3$ , $Re_\infty /{\rm cm} = 3.3 \times 10^4$ . Air was used as the working gas for both the freestream and the jet. Schlieren pictures were used for the visualisation of the three-dimensional structures around the jet. Pressure, normal force and pitching moment measurements were conducted to quantitatively study the interaction region and its effects on the vehicle. An analytical algorithm has been developed to predict the shape of the separation front around the body.

Navier-Stokes Computations of the NREL Airfoil Using a κ — ω Turbulent Model at High Angles of Attack

Abstract: This paper presents a two-dimensional numerical simulation of the turbulent flow fields for the NREL (National Renewable Energy Laboratory) S809 airfoil. The flow is modeled as steady, viscous, turbulent, and incompressible. The pseudo-compressible formulation is used for the time-averaged Navier-Stokes equations so that a time marching scheme developed for the compressible flow can be applied directly. The turbulent flow is simulated using Wilcox`s modified {kappa}-{omega} model to account for the low Reynolds number effects near a solid wall and the model`s sensitivity to the freestream conditions. The governing equations are solved by an implicit approximate-factorization scheme. To correctly model the convection terms in the mean-flow and turbulence model equations, the symmetric TVD (Total Variational Diminishing) scheme is incorporated. The methodology developed is then applied to analyze the NREL S809 airfoil at various angles of attack ({alpha}) from 1 to 45 degrees. The accuracy of the numerical results is compared with the available Delft wind tunnel test data. For comparison, two Eppler code results at low angles of attack are also included. Depending on the value of {alpha}, preliminary results show excellent to fairly good agreement with the experimental data. Directions for future work are also discussed.

Study of subsonic base cavity flowfield structure using particle image velocimetry

Abstract: A new particle image velocimetry system has been used to study the near-wake structure of a two-dimensional base in subsonic flow to determine the fluid dynamic mechanisms of observed base drag reduction in the presence of a base cavity. Experiments were done over a range of freestream Mach numbers up to 0.8, including local flowfield velocities over 300 m/s. Effects of the base cavity on the von Karman vortex street wake were found to be related to the expansion and diffusion of vortices near the cavity, although the effects are of small magnitude and no significant change in the vortex formation location or path was observed. The base cavity effects are also less significant at higher freestream velocities due to the formation of vortices further downstream from the base. The base cavity drag reduction was found to be mainly due to the displacement of the base surface to a location upstream of the low-pressure wake vortices, with only a slight modification in the vortex street itself.

Measurements of compressible turbulent flow structure in a supersonic mixing layer

Abstract: Multiple overheat cross-wire, parallel hot-wire, shadowgraph image processing and high-frequency response pitot probe surveys were taken in a two-dimensional, adiabatic, supersonic, free mixing layer, consisting of Mach 1.8 air (Re = 7 x 10 6 /m) injected tangentially into a Mach 4.0 freestream (Re = 67 x 10 6 /m). Mass flux turbulence intensity profiles in three dimensions were acquired, and the axial component was found to be about 25% larger than the transverse and spanwise components. Mean Favre-averaged velocity fluctuation data indicated that the axial component was negative, while both the transverse and span were positive. Turbulent flow structure angles and length scales were measured in the fully developed region of the flow. The structure angles were found to be consistent with previous compressible boundary layer results ; however, the magnitude here was found to decrease across the mixing layer. Mach number fluctuation levels were also measured. The peak Mach number fluctuation level in this flow was about 10%.

The Receptivity of Flat-Plate Boundary-Layers with Two-Dimensional Roughness Elements to Freestream Sound and Its Control

Abstract: The receptivity of a flat-plate boundary layer to freestream sound when one or two two-dimensional roughness elements are attached to the surface is experimentally investigated. It is shown that a thin two-dimensional roughness element together with sound can generate T-S wave and its amplitude can be controlled by adding the second roughness element and changing the distance between the two elements. This effect is explained by the superposition of the waves generated by each roughness element. It is also shown that in the case of single roughness element, the amplitude of the generated T-S wave is a function of roughness-width. It is revealed that the T-S wave generated at the leading edge of the roughness element and that generated at the trailing edge are 180 degrees out of phase and they either cancel or amplify depending on the distance between the edges.

Computed Effects of Heat Transfer on the Transonic Flow over an Aerofoil

Abstract: The effects of the model surface to freestream adiabatic temperature ratio (Tw/Ta^) on subsonic flows at zero pressure gradient and transonic flow over a NACA0012 aerofoil are evaluated using a computational fluid dynamic approach. The analysis, based on the thin layer Navier-Stokes equations with a Bald win-Lorn ax turbulence model, indicated that surface heat transfer has significant effects on both subsonic and transonic flows confirming some of the experimental data available. The results have implications in wind-tunnel testing at nonadiabatic surface conditions.

Simulated Rarefied Entry of the Galileo Probe into the Jovian Atmosphere

Abstract: Flowfield properties and aerodynamics are computed with a direct simulation Monte Carlo method in the rarefied flow regime during entry of the Galileo Probe into the atmosphere of Jupiter. The objective is to predict accurately the vehicle's drag coefficient, which is needed to assess atmospheric properties from the onboard atmospheric structure experiment, where highly sensitive accelerometers will measure the drag effects to within 10~5 m/s2 during the initial entry phase at high altitudes. The corresponding flow rarefaction extends from the free-molecule limit to the near-continuum transition regime (Re^ < 1000). Simulation results, employing a simple radiative equilibrium surface model, indicate that CD varies from 2.1 at the free-molecule limit down to 1.6 at Re^ = 1000. Results compared very well to those from ballistic-range experiments. Detailed material response of the carbonphenolic heat shield was then coupled directly into the DSMC code to account accurately for conductivity, heat capacity, and pyrolysis, and the simulations were repeated. The predicted pyrolysis mass efflux was 8-14 times higher than the incident freestream mass flux and had significant effects on the drag.

Experimental Study of the Unsteady Boundary Layer Behavior on a Turbine Cascade

Abstract: The unsteady boundary layer behavior on a turbine cascade is experimentally investigated and the results are presented in this paper. To perform a detailed study on unsteady cascade aerodynamics and heat transfer, a new large-scale, high-subsonic research facility for simulating the periodic unsteady flow has been developed. It is capable of sequentially generating up to four different unsteady inlet flow conditions that lead to four different passing frequencies, wake structures, and freestream turbulence intensities. For a given Reynolds number, three different unsteady wake formations are utilized. Detailed unsteady boundary layer velocity, turbulence intensity, and pressure measurements are performed along the suction and pressure surfaces of one blade. The results presented in the temporal-spatial domain display the transition and further development of the boundary layer, specifically the ensemble-averaged velocity and turbulence intensity.Copyright © 1995 by ASME

Tunable diode laser measurements on nitric oxide in a hypersonic wind tunnel

Abstract: A tunable infrared diode laser was used in the F4 high-enthalpy hypersonic wind tunnel at ONERA to make line-of-sight measurements on nitric oxide (NO). The NO is generated in the arc chamber and remains frozen in the hypersonic flow. We report the results of a medium enthalpy run. The Ω (14.5), Ω = 3/2 doublet was recorded simultaneously in two views : one normal to the flow and the other biased at 63-deg to the flow axis. Spectra were recorded at 2 ms intervals during a tunnel run, and 30 sets of spectra were analyzed by a least-squares fitting procedure. The mean of the velocity measurements inferred from the Doppler shift of the doublet was 2740 m/s (±5%). The velocity decayed about 10% over 60 ms. The temperature and NO column density determined by the least-squares fits were subject to much larger uncertainties. The static temperature of the freestream is estimated to be 168 K (±35%). The median inferred NO column density of 4 X 10 17 molecule/cm 2 corresponds to an NO number density of order 5 x 10 15 molecules/cm 3 in the freestream.

Interpretation of experimental high-alpha aerodynamics. Implications for flight prediction

Abstract: Nomenclature b = wingspan, or largest dimension C — aerodynamic coefficient; with no superscript, in body axes system Cij = dCi/d(jl/2V), i = /, m, n\ j = q, q', d, a(I = c)'J =p,j8, (/ = b) Cik = dCf/dk, i = l,m,n\k = a, /3, a Cl = rolling moment coefficient Cm = pitching moment coefficient C,, = yawing moment coefficient Cp = static pressure coefficient CY = side force coefficient c = mean aerodynamic chord d — body maximum diameter h = height of test section h() = model length / = generalized reference length M^ = freestream Mach number /?, px = local static pressure /?, q, r = body-axes angular velocities q^ = freestream dynamic pressure Re, Reynolds number based on /, / c, b, or d S = reference area V, V^ = freestream velocity w = minimum dimension of test section X, Y, Z = inertial axes system Xf, Yf; Zf = flight coordinate system, Fig. 2 x, _y , z = body axes system a, /3 = angles of attack and sideslip A = increment or amplitude A = inclination of rotation axis, Fig. 6 = coning rate, parameter fib/2V or £1 a) = reduced circular frequency, coll(2V), where / = c or b, as appropriate

A driver gas detection device for shock tunnels

Abstract: A device has been produced which can detect the contamination of the test gas by the driver gas in a reflected shock tunnel. This device monitors the static pressure in a converging duct. The duct is designed to choke at a predetermined contamination level due to the change in the specific heat ratio produced by the contaminants. Experimental results are given for a freestream enthalpy of nominally 6 MJ/kg.