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Pressure distribution data at subsonic and transonic speeds for calibrating hypersonic conical pressure probe for X-15 aircraft
The changes in aerodynamic characteristics due to real-gas effects associated with high speed flight (characterized by large shock density ratios) are primarily the result of changes in surface pressures acting on the forebody. The surface pressures are affected by a change in shock density ratio (real-gas effects) in two ways. First, the level of pressure at the stagnation point relative to freestream dynamic pressure is changed, and second, the distribution of surface pressure relative to stagnation-point pressure is changed. The density-ratio effect on the stagnation point pressure level can be estimated by considering the flow of a perfect gas about a blunt body.
The Aerothermal Design Space Interpolation (ADSI) tool is used to interpolate databases of previously computed computational fluid dynamic solutions for test articles in a NASA Ames arc jet facility. The arc jet databases are generated using an Navier-Stokes flow solver using previously determined best practices. The arc jet mass flow rates and arc currents used to discretize the database are chosen to span the operating conditions possible in the arc jet, and are based on previous arc jet experimental conditions where possible. The ADSI code is a database interpolation, manipulation, and examination tool that can be used to estimate the stagnation point pressure and heating rate for user-specified values of arc jet mass flow rate and arc current. The interpolation is performed in the other direction (predicting mass flow and current to achieve a desired stagnation point pressure and heating rate). ADSI is also used to generate 2-D response surfaces of stagnation point pressure and heating rate as a function of mass flow rate and arc current (or vice versa). Arc jet test data is used to assess the predictive capability of the ADSI code.
A combination probe for time-resolved measurements of unsteady compressible flows in transonic wind tunnels is described. The probe measures stagnation (total) temperature and pressure, static pressure, and flow angles in two planes. From these, the fluctuating mass flux, Mach number, and velocity as well as their components in three directions can be deduced. The combination probe consists of a dual hot-wire aspirating temperature and pressure probe mounted piggyback with a high-frequency angle probe. The angle probe has four surface-mounted silicon pressure sensors. A scheme is described for retrieving from the four pressure signals the stagnation and static pressures, Mach number, and flow angles in two planes. The calibrations forming the base for this procedure, obtained from steady-state tests, are given. Typical data obtained in the Karman vortex street shed from a cylinder and at the exit of a Mach 0.4 air jet are presented.
A combination probe for time-resolved measurements for unsteady compressible flows is described. The probe measures stagnation (total) temperature and pressure, static pressure, and flow angles in two planes. From these, the fluctuating mass flux, Mach number, and velocity as well as their components in three directions can be deduced. The combination probe consists of a dual hot-wire aspirating temperature and pressure probe mounted piggyback with a high-frequency angle probe. The angle probe has four surface-mounted silicon pressure sensors. A scheme is described for retrieving from the four pressure signals the stagnation and static pressures, Mach number, and flow angles in two planes. The calibrations forming the basis for this procedure, obtained from steady-state tests, are given. Typical data obtained in the Karman vortex street shed from a cylinder and at the exit of a Mach-0.4 air jet are presented.
A newly developed, 3-mm-diam, dual hot-wire aspirating probe was used to measure the time-resolved stagnation temperature and pressure in a transonic cryogenic wind tunnel. Measurements were taken in the freestream of the settling chamber and test section. Data were also obtained in the unsteady wake shed from an airfoil oscillating at 5 Hz. The investigation revealed the presence of large fluctuations in the settling chamber occuring at the blade passing frequency of the driving fan of the tunnel. These fluctuations decrease at the test section. The rms value of the fluctuating stagnation pressure decreased from 17.5 percent in the settling chamber to 3.7 percent in the test section. Fluctuating stagnation temperature decreased from 12.3 percent to 8.4 percent. Measurements in the wake of the oscillating airfoil showed a fluctuating stagnation temperature of as much as 42 K in rms value.
The Hypersonic Materials Environmental Test System (HyMETS) arc-jet facility located at the NASA Langley Research Center in Hampton, Virginia, is primarily used for the research, development, and evaluation of high-temperature thermal protection systems for hypersonic vehicles and reentry systems. In order to improve testing capabilities and knowledge of the test article environment, an effort is underway to computationally simulate the flow-field using computational fluid dynamics (CFD). A detailed three-dimensional model of the arc-jet nozzle and free-jet portion of the flow-field has been developed and compared to calibration probe Pitot pressure and stagnation-point heat flux for three test conditions at low, medium, and high enthalpy. The CFD model takes into account uniform pressure and non-uniform enthalpy profiles at the nozzle inlet as well as catalytic recombination efficiency effects at the probe surface. Comparing the CFD results and test data indicates an effectively fully-catalytic copper surface on the heat flux probe of about 10% efficiency and a 2-3 kpa pressure drop from the arc heater bore, where the pressure is measured, to the plenum section, prior to the nozzle. With these assumptions, the CFD results are well within the uncertainty of the stagnation pressure and heat flux measurements. The conditions at the nozzle exit were also compared with radial and axial velocimetry. This simulation capability will be used to evaluate various three-dimensional models that are tested in the HyMETS facility. An end-to-end aerothermal and thermal simulation of HyMETS test articles will follow this work to provide a better understanding of the test environment, test results, and to aid in test planning. Additional flow-field diagnostic measurements will also be considered to improve the modeling capability.
A newly developed, 3-mm-diam, dual hot-wire aspirating probe was used to measure the time-resolved stagnation temperature and pressure in a transonic cryogenic wind tunnel. The probe consists of two coplanar constant temperature hot wires at different overheat ratios operating in a 1.5-mm-diam channel with a choked exit. Thus, the constant Mach number flow by the wires is influenced only by free-stream stagnation temperature and pressure. Diffusion of the free-stream Mach number to a lower value in the channel reduces the dynamic drag on the hot-wire. Frequency response of the present design is dc to 20 kHz. The probe was used to measure the unsteady wake shed from an oscillating airfoil tested in the 0.3-m Transonic Cryogenic Tunnel at NASA-Langley Research Center. The hot-wire lasted for more than ten hours before breaking, proving the ruggedness of the probe and the usefulness of the technique in a high dynamic pressure, transonic cryogenic wind tunnel. Typical data obtained from the experiment are presented after reduction to stagnation pressure and temperature.
Low-pressure blowoff experiments were conducted with a stagnation flame stabilized on the forward tip of cast PMMA rods in a vertical wind tunnel. Pressure, forced flow velocity, gravity, and ambient oxygen concentration were varied. Stagnation flame blowoff is determined from a time-stamped video recording of the test. The blowoff pressure is determined from test section pressure transducer data that is synchronized with the time stamp. The forced flow velocity is also determined from the choked flow orifice pressure. Most of the tests were performed in normal gravity, but a handful of microgravity tests were also conducted to determine the influence of buoyant flow velocity on the blowoff limits. The blowoff limits are found to have a linear dependence between the partial pressure of oxygen and the total pressure, regardless of forced flow velocity and gravity level. The flow velocity (forced and/or buoyant) affects the blowoff pressure through the critical Damkohler number residence time, which dictates the partial pressure of oxygen at blowoff. This is because the critical stretch rate increases linearly with increasing pressure at low pressure (sub-atmospheric pressures) since a second-order overall reaction rate with two-body reactions dominates in this pressure range.
A thermochemically-closed sonic-flow inversion framework (TSIF) is developed to infer bulk enthalpy and total temperature upstream of a choked nozzle in arc-jet flows. The formulation recasts a pressure-rise total enthalpy quantification technique as an inverse problem in characteristic-velocity c * space using measured mass flow rate, upstream total pressure, gas composition, and nozzle throat geometry as inputs. Unlike calorimetric energy-balance approaches or optical diagnostics, the method relies primarily on routinely measured facility quantities combined with explicit thermochemical closure. Thermochemical states are obtained using NASA’s open-source Chemical Equilibrium with Applications (CEA) code, enabling construction of a chemistry-consistent relation between characteristic velocity, total enthalpy, and total temperature under equilibrium or frozen assumptions. A discharge coefficient is self-calibrated using cold-flow (arc-off) operation data and applied to hot-flow (arc-on) measurements, enabling upstream losses to be accounted for without empirical correlations. The framework is applied to air, N 2 , and CO 2 –N 2 arc-jet flows and demonstrates expected trends for the inferred thermochemical states as function of arc power, specific energy input, mass-flow, heater configuration, and test gas. In the air limit, under equilibrium assumptions, the method recovers the classical high-enthalpy asymptotic correlation of Winovich with a mean residual of 4.4%, demonstrating compatibility with established sonic-flow scaling, while extending applicability to arbitrary multi-species mixtures and non-equilibrium chemistry. The framework provides a mixture-flexible methodology for determining bulk thermochemical states in modern arc-jet environments using routine facility pressure, mass-flow, gas-composition, and nozzle-geometry information together with a cold-flow calibration.
Two circular conical configurations having 15 deg half-angles were tested in laminar boundary layer at a Mach number of 6 and angles of attack up to 90 deg. One cone had a sharp nose and a fineness ratio of 1.87 and the other had a spherically blunted nose with a bluntness ratio of 0.1428 and a fineness ratio of 1.66. Pressure measurements and schlieren pictures of the flow showed that near-conical flow existed up to an angle of attack of approximately 60 deg. At angles of attack above 70 deg high-pressure areas were present near the base and the bow shock wave was considerably curved. Comparison of the results with simply applied theories showed that on the stagnation line pressures may be predicted by Newtonian theory, and heat transfer by local yawed-cylinder theory based on the yaw angle of the windward generator and the local radius of the cone. Base effects increased the heat transfer in a region extending forward approximately 15 to 30 percent of the windward generator. Circumferential pressure distributions were higher than the corresponding Newtonian distribution and a better prediction was obtained by modifying the theory to match the pressure at 90 deg from the windward generator to that on the surface of the cone at an angle of attack of 0 deg. Circumferential heat-transfer distributions were predicted satisfactorily up to about 60 deg from the stagnation line by using Lees' heat-flux distribution based on the Newtonian pressure. The effects of nose bluntness at large angles of attack were very small in the region beyond two nose radii from the point of tangency.
A high-frequency combination probe was used to measure dynamic flow quality in the test section of the NASA Langley 0.3-m Transonic Cryogenic Tunnel. The probe measures fluctuating stagnation (total) temperature and pressure, static pressure, and flow angles in two orthogonal planes. Simultaneous measurements of unsteady total temperature and pressure were also made in the settling chamber of the tunnel. The data show that the stagnation temperature fluctuations remain constant, and the stagnation pressure fluctuations increase by a factor of two, as the flow accelerates from the settling chamber to the test section. In the test section, the maximum rms value of the normalized fluctuating velocity is 0.7 percent. Correlation coefficients failed to show vorticity, entropy, or sound as the dominant mode of turbulence in the tunnel.
We present here the detailed measurements of radial distribution of the magnetic field in a gas-puff z-pinch plasma at the final stages of the implosion phase and at stagnation. While the measurements are chordal, the radial distribution of different charge states was utilized to measure the magnetic field locally for certain radii, so that unlike chordal measurements in general, the magnetic field radial distribution was obtained with no need for the Abel inversion of the data. The distribution was measured using the Zeeman effect via a novel spectroscopic technique, at several axial locations, and demonstrates striking features such as the peak field remaining at a radius much larger than the stagnation radius at all times. Furthermore, while the distribution observed is sometimes monotonic with respect to the radius, it is often not, a behavior that can be linked to 2D features in the plasma column resulting from the Rayleigh–Taylor instability. The current flowing through the stagnating plasma was found to be a small fraction of the total current, resulting in clearly insufficient magnetic pressure to balance the plasma pressure at stagnation. The magnetic field data, taken over several axial positions, are used to obtain the true inductance in the imploding plasma for the first time; it is found that the data cannot explain the current turnover at stagnation. A simulation with the MACH2-Tabular Collisional-Radiative Equilibrium magnetohydrodynamics code in the r–z plane shows that the peak of the magnetic field pinches to a much smaller radius than is observed in the spectroscopic data. Furthermore, the depth of the computed current turnover at stagnation is smaller than the measured one. The two observed features of a radially extended magnetic field at stagnation together with a deep current turnover are a challenge to match in simulations. Various calculations and estimates of the inductive and resistive load voltages are examined to ascertain if they are responsible for the observed current notch. Lastly, the results demonstrate that the knowledge of the true inductance in the driven load requires such magnetic-field-distribution measurements and that imaging data or electrical measurements are insufficient.
An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to examine the influence of arrays of regularly spaced roughness elements on instability growth and transition. The primary element shape was a pair of elliptical planform ramps that were inclined at equal and opposite angles with respect to the local streamwise direction. The element shapes were designed to induce transient growth disturbances that would lead to sustained azimuthal modulation of the boundary layer flow while limiting the nearfield disturbances to avoid an immediate, i.e., effective tripping of the boundary layer. The bulk of the run matrix consisted of testing different element height sat free stream unit Reynolds numbers ranging from 9.8 to 13.1 million per meter. Other element shapes previously designed for tripping hypersonic boundary layers were also implemented. The model was instrumented with surface mounted Kulite ® and PCB ® pressure transducers and thermocouples. Spectra from the PCBs ® indicated clear suppression of the second-mode instability; however, neither the PCB ® spectra nor the heat transfer data presented strong evidence for delayed turbulent flow. Complementary stability computations likewise demonstrated second-mode reduction, particularly just downstream of the roughness, but also revealed a rise in first mode (streak-instability) amplitudes from the baseline that was likely responsible for the earlier transition observed for taller roughness cases.
An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to examine the influence of arrays of regularly spaced roughness elements on instability growth and transition. The primary element shape was a pair of elliptical planform ramps that were inclined at equal and opposite angles with respect to the local streamwise direction. The element shapes were designed to induce transient growth disturbances that would lead to sustained azimuthal modulation of the boundary layer flow while limiting the nearfield disturbances to avoid an immediate, i.e., effective tripping of the boundary layer. The bulk of the run matrix consisted of testing different element height sat free stream unit Reynolds numbers ranging from 9.8 to 13.1 million per meter. Other element shapes previously designed for tripping hypersonic boundary layers were also implemented. The model was instrumented with surface mounted Kulite ® and PCB ® pressure transducers and thermocouples. Spectra from the PCBs ® indicated clear suppression of the second-mode instability; however, neither the PCB ® spectra nor the heat transfer data presented strong evidence for delayed turbulent flow. Complementary stability computations likewise demonstrated second-mode reduction, particularly just downstream of the roughness, but also revealed a rise in first mode (streak-instability) amplitudes from the baseline that was likely responsible for the earlier transition observed for taller roughness cases.
An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to examine the influence of arrays of regularly spaced roughness elements on instability growth and transition. The primary element shape was a pair of elliptical planform ramps that were inclined at equal and opposite angles with respect to the local streamwise direction. The element shapes were designed to induce transient growth disturbances that would lead to sustained azimuthal modulation of the boundary layer flow while limiting the nearfield disturbances to avoid an immediate, i.e., effective tripping of the boundary layer. The bulk of the run matrix consisted of testing different element height sat free stream unit Reynolds numbers ranging from 9.8 to 13.1 million per meter. Other element shapes previously designed for tripping hypersonic boundary layers were also implemented. The model was instrumented with surface mounted Kulite ® and PCB ® pressure transducers and thermocouples. Spectra from the PCBs ® indicated clear suppression of the second-mode instability; however, neither the PCB ® spectra nor the heat transfer data presented strong evidence for delayed turbulent flow. Complementary stability computations likewise demonstrated second-mode reduction, particularly just downstream of the roughness, but also revealed a rise in first mode (streak-instability) amplitudes from the baseline that was likely responsible for the earlier transition observed for taller roughness cases.
The SEADS pressure data obtained from the Shuttle flight 61-C are analyzed in conjunction with the preflight database. Based on wind tunnel data, the sensitivity of the Shuttle Orbiter stagnation region pressure distribution to angle of attack and Mach number is demonstrated. Comparisons are made between flight and wind tunnel SEADS orifice pressure distributions at several points throughout the re-entry. It is concluded that modified Newtonian theory provides a good tool for the design of a flush air data system, furnishing data for determining orifice locations and transducer sizing. Ground-based wind tunnel facilities are capable of providing the correction factors necessary for the derivation of accurate air data parameters from pressure data.
A thermochemically-closed sonic-flow inversion framework (TSIF) is developed to infer bulk enthalpy and total temperature upstream of a choked nozzle in arc-jet flows. The formulation recasts a pressure-rise total enthalpy quantification technique as an inverse problem in characteristic-velocity c * space using measured mass flow rate, upstream total pressure, gas composition, and nozzle throat geometry as inputs. Unlike calorimetric energy-balance approaches or optical diagnostics, the method relies primarily on routinely measured facility quantities combined with explicit thermochemical closure. Thermochemical states are obtained using NASA’s open-source Chemical Equilibrium with Applications (CEA) code, enabling construction of a chemistry-consistent relation between characteristic velocity, total enthalpy, and total temperature under equilibrium or frozen assumptions. A discharge coefficient is self-calibrated using cold-flow (arc-off) operation data and applied to hot-flow (arc-on) measurements, enabling upstream losses to be accounted for without empirical correlations. The framework is applied to air, N 2 , and CO 2 –N 2 arc-jet flows and demonstrates expected trends for the inferred thermochemical states as function of arc power, specific energy input, mass-flow, heater configuration, and test gas. In the air limit, under equilibrium assumptions, the method recovers the classical high-enthalpy asymptotic correlation of Winovich with a mean residual of 4.4%, demonstrating compatibility with established sonic-flow scaling, while extending applicability to arbitrary multi-species mixtures and non-equilibrium chemistry. The framework provides a mixture-flexible methodology for determining bulk thermochemical states in modern arc-jet environments using routine facility pressure, mass-flow, gas-composition, and nozzle-geometry information together with a cold-flow calibration.