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It is shown that the external vorticity condition of Hayes is generally limited to the stagnation point of axisymmetric blunt bodies because of compatibility with the boundary layer momentum equation. It is also shown that in the stagnation region, the condition is not compatible with the cold wall approximation of Lees.
NASA's Genesis sample return mission will be the first to return material from beyond the Earth-Moon system. NASA Langley Research Center supported this mission with aerothermodynamic analyses of the sample return capsule. This paper provides an overview of that effort. The capsule is attached through its forebody to the spacecraft bus. When the attachment is severed prior to Earth entry, forebody cavities remain. The presence of these cavities could dramatically increase the heating environment in their vicinity and downstream. A combination of computational fluid dynamics calculations and wind tunnel phosphor thermography tests were employed to address this issue. These results quantify the heating environment in and around the cavities, and were a factor in the decision to switch forebody heat shield materials. A transition map is developed which predicts that the flow aft of the penetrations will still be laminar at the peak heating point of the trajectory. As the vehicle continues along the trajectory to the peak dynamic pressure point, fully turbulent flow aft of the penetrations could occur. The integrated heat load calculations show that a heat shield sized to the stagnation point levels will be adequate for the predicted environment aft of the penetrations.
With the development of new manned spaceflight capabilities including NASA's Orion capsule and the Space-X Dragon capsule, there is a renewed importance of understanding the dynamics of ablative thermal protection systems. To this end, a new inductively coupled plasma torch facility is being developed at UT-Austin. The torch operates on argon and/or air at plasma powers up to 50 kW. In the present configuration the flow issues from a low-speed subsonic nozzle and the hot plume is characterized using slug calorimetry and emission spectroscopy. Preliminary measurements using emission spectroscopy have indicated that the torch is capable of producing an air plasma with a temperature between 6,000 K and 8,000 K depending on the power and flow settings and an argon plasma with a temperature of approximately 12,000 K. The operation envelope was measured, and heat flux measured for every point within the envelope using both a slug calorimeter and a Gardon gauge heat flux sensor. The torch was found to induce a stagnation point heat flux of between 90 and 225 W/sq cm.
The mARC II is a 30 kW arc-jet facility at NASA Ames Research Center developed to produce high-enthalpy flows for low-cost technology development purposes. In this work, we introduce new measurement capabilities following the latest facility upgrades and begin characterizing the operating envelope for the two-disk arc-heater configuration with air as the working gas. The lower bound of the envelope corresponds to the lowest set current and the farthest sensor distance from the nozzle (𝐼set = 40 A, 𝑧 = 69 mm), while the upper bound corresponds to the highest current and closest sensor distance (𝐼set = 200 A, 𝑧 = 2 mm). Stagnation point heat flux was measured using a water-cooled Gardon gauge (⌀4.76 mm, 3/16" hemispherical), with values ranging from 15 ≤ q̇₀ ≤ 900 W/cm2. Stagnation pressure was measured using a water-cooled Pitot probe (⌀4.76 mm, 3/16" hemispherical), with values ranging from 18 ≤ p₀ ≤ 1100 Pa. The upgraded vacuum system is demonstrated to significantly extend the lower end of the operating envelope of the mARC II. Additionally, measurements of bulk enthalpy (3 ≤ h ≤ 13 MJ/kg) and stagnation enthalpy (4 ≤ h₀ ≤ 34 MJ/kg) demonstrate the mARC II facility's high-enthalpy capabilities. Completion of the operating envelope characterization is ongoing in tandem with the development of complementary diagnostic capabilities and numerical simulations.
The mARC II is a 30 kW arc-jet facility at NASA Ames Research Center developed to produce high-enthalpy flows for low-cost technology development purposes. In this work, we introduce new measurement capabilities following the latest facility upgrades and begin characterizing the operating envelope for the two-disk arc-heater configuration with air as the working gas. The lower bound of the envelope corresponds to the lowest set current and the farthest sensor distance from the nozzle (𝐼set = 40 A, 𝑧 = 69 mm), while the upper bound corresponds to the highest current and closest sensor distance (𝐼set = 200 A, 𝑧 = 2 mm). Stagnation point heat flux was measured using a water-cooled Gardon gauge (⌀4.76 mm, 3/16" hemispherical), with values ranging from 15 ≤ q̇₀ ≤ 900 W/cm 2 . Stagnation pressure was measured using a water-cooled Pitot probe (⌀4.76 mm, 3/16" hemispherical), with values ranging from 18 ≤ p₀ ≤ 1100 Pa. The upgraded vacuum system is demonstrated to significantly extend the lower end of the operating envelope of the mARC II. Additionally, measurements of bulk enthalpy (3 ≤ h ≤ 13 MJ/kg) and stagnation enthalpy (4 ≤ h₀ ≤ 34 MJ/kg) demonstrate the mARC II facility's high-enthalpy capabilities. Completion of the operating envelope characterization is ongoing in tandem with the development of complementary diagnostic capabilities and numerical simulations.
A model magnetosphere for Mercury is presented using an upstream image-dipole and nightside 2-dimensional tail current sheet method. The tail field is represented by an analytical formulation. Magnetic field data from the Mercury 1 encounter by Mariner 10 in March 1974 are used to determine quantitative parameters of the model magnetosphere, using the method of least squares. The magnetopause crossing points directly observed are used to determine the size of the magnetosphere, and the solar wind conditions are used to determine the magnetospheric field at the stagnation point. The model produces a magnetosphere-like region with planetary field lines that are confined in nearly circular cross-sections transverse to the sun-planet line. Results are used to show geometry, field line configuration, and contours of constant field intensity inside the magnetosphere.
A wind tunnel test program has been conducted to define convective heating environments on the back-face of a Hypersonic Inflatable Aerodynamic Decelerator aeroshell. Wind tunnel testing was conducted at Mach 6 and Mach 10 at unit Reynolds numbers from 0.5×10(exp 6)/ft to 3.9×10(exp 6)/ft on a 6.3088 in diameter aeroshell model. Global heating data were obtained through phosphor thermography on the aeroshell back face, as well as on the payload and the aeroshell front face. For all test conditions, laminar flow was produced on the aeroshell front face, while the separated wake shear layer and aeroshell back-face boundary layer were transitional or turbulent. Along the leeward centerline of the aeroshell back face and payload centerbody, heating levels increased with both free stream Reynolds number and angle of attack. The Reynolds number dependency was due to increasing strength of wake turbulence with Reynolds number. The angle-of-attack dependency was due to movement of the wake-vortex reattachment point on the aeroshell back face. The maximum heating levels on the aeroshell back face and payload were approximately 5% to 6%, respectively, of the aeroshell front-face stagnation point. To allow for extrapolation of the ground test data to flight conditions, the back face and payload heating levels were correlated as a function of aeroshell front-face peak momentum thickness Reynolds numbers.
An investigation was made in the NACA 7- by 10-foot wind tunnel of a large-chord wing model with a duct to house a simulated radiator suitable for a liquid-cooled engine. The duct was expanded to reduce the radiator losses, and the installation of the duct and radiator was made entirely within the wing to reduce form and interference drag. The tests were made using a two-dimensional flow set-up with a full-span duct and radiator. Section aerodynamic characteristics of the basic airfoil are given and also curves showing the characteristics of the various duct-radiator combinations. An expression for efficiency, the primary criterion of merit of any duct, and the effect of the several design parameters of the duct-radiator arrangement are discussed. The problem of throttling is considered and a discussion of the power required for cooling is included. It was found that radiators could be mounted in the wing and efficiently pass enough air for cooling with duct outlets located at any point from 0.25c to 0.70c from the wing leading edge on the upper surface. The duct-inlet position was found to be critical and, for maximum efficiency, had to be at the stagnation point of the airfoil and to change with flight attitude. The flow could be efficiently throttled only by a simultaneous variation of duct inlet and outlet sizes and of inlet position. It was desirable to round both inlet and outlet lips. With certain arrangements of duct, the power required for cooling at high speed was a very low percentage of the engine power.
The objective of the proposed research project was the analysis of single point closures based on probability density function (pdf) and characteristic functions and the development of a prediction method for the joint velocity-scalar pdf in turbulent reacting flows. Turbulent flows of boundary layer type and stagnation point flows with and without chemical reactions were be calculated as principal applications. Pdf methods for compressible reacting flows were developed and tested in comparison with available experimental data. The research work carried in this project was concentrated on the closure of pdf equations for incompressible and compressible turbulent flows with and without chemical reactions.
Numerical solutions to the Euler equations for transonic flow over a circular cylinder suggest that the inviscid flow separates ahead of the rear stagnation point. The present understanding of this phenomenon and various solutions presented at a workshop on this subject are discussed. All the evidence suggests that inviscid separation is a feature of the Euler equations induced by vorticity and/or stagnation pressure loss in the flow. The possibility of multiple solutions suggested by Hafez (1981) and the oscillatory behavior for the full circle observed by Steger (1977) are seen as needing further investigation.
The degree of compression sustained by the solar wind field as it convects to the magnetopause has been determined empirically with magnetometer data from ISEE 3 in the solar wind and ISEE 1 in the magnetosheath. At the stagnation point, the strength B(SH) of the magnetosheath field may be expressed as B(SH) = square root of 4 X B(SW)' X B(ST), which implies that B(SH) is equal to the geometric mean of the stagnation field strength B(ST) and the shocked solar wind field, approximated by 4 X B(SW)'. For the usual spiral wind pattern of the solar wind field, B(SH) on the duskside of the magnetopause is about 20 percent stronger than it is on the downside. No dependence of B(SH) on the sign of the north-south component of the solar wind field is apparent.
A formulation of finite-rate ablation surface boundary conditions, including oxidation, nitridation, and sublimation of carbonaceous material with pyrolysis gas injection, has been developed based on surface species mass conservation. These surface boundary conditions are discretized and integrated with a Navier-Stokes solver. This numerical procedure can predict aerothermal heating, chemical species concentration, and carbonaceous material ablation rate over the heatshield surface of re-entry space vehicles. In this study, the gas-gas and gas-surface interactions are established for air flow over a carbon-phenolic heatshield. Two finite-rate gas-surface interaction models are considered in the present study. The first model is based on the work of Park, and the second model includes the kinetics suggested by Zhluktov and Abe. Nineteen gas phase chemical reactions and four gas-surface interactions are considered in the present model. There is a total of fourteen gas phase chemical species, including five species for air and nine species for ablation products. Three test cases are studied in this paper. The first case is a graphite test model in the arc-jet stream; the second is a light weight Phenolic Impregnated Carbon Ablator at the Stardust re-entry peak heating conditions, and the third is a fully dense carbon-phenolic heatshield at the peak heating point of a proposed Mars Sample Return Earth Entry Vehicle. Predictions based on both finite-rate gas- surface interaction models are compared with those obtained using B' tables, which were created based on the chemical equilibrium assumption. Stagnation point convective heat fluxes predicted using Park's finite-rate model are far below those obtained from chemical equilibrium B' tables and Zhluktov's model. Recession predictions from Zhluktov's model are generally lower than those obtained from Park's model and chemical equilibrium B' tables. The effect of species mass diffusion on predicted ablation rate is also examined.
An analytical model for separated airfoil flows is presented which is based on experimentally observed physical phenomena. These include a free stagnation point aft of the airfoil and a standing vortex in the separated region. A computer program is described which iteratively matches the outer potential flow, the airfoil turbulent boundary layer, the separated jet entrainment, mass conservation in the separated bubble, and the rear stagnation pressure. Separation location and pressure are not specified a priori. Results are presented for surface pressure coefficient and compared with experiment for three angles of attack for a GA(W)-1, 17% thick airfoil.
Heat-transfer coefficients and pressure distributions were obtained on a 4-inch-diameter flat-face cylinder in the Langley Unitary Plan wind tunnel. The measured stagnation heat-transfer coefficient agrees well with 55 percent of the theoretical value predicted by the modified Sibulkin method for a hemisphere. Pressure measurements indicated the dimensionless velocity gradient parameter r du\ a(sub t) dx, where x=0 at the stagnation point was approximately 0.3 and invariant throughout the Mach number range from 2.49 to 4.44 and the Reynolds number range from 0.77 x 10(exp 6) to 1.46 x 10(exp 6). The heat-transfer coefficients on the cylindrical afterbody could be predicted with reasonable accuracy by flat-plate theory at an angle of attack of 0 deg. At angles of attack the cylindrical afterbody stagnation-line heat transfer could be computed from swept-cylinder theory for large distances back of the nose when the Reynolds number is based on the distance from the flow reattachment points.
The heat transfer behavior of accreting ice surfaces in natural (flight test) and simulated (wind tunnel) cloud icing conditions were studied. Observations of wet and dry ice growth regimes as measured by ultrasonic pulse echo techniques were made. Observed wet and dry ice growth regimes at the stagnation point of a cylinder were compared with those predicted using a quasi steady state heat balance model. A series of heat transfer coefficients were employed by the model to infer the local heat transfer behavior of the actual ice surfaces. The heat transfer in the stagnation region was generally inferred to be higher in wind tunnel icing tests than in natural flight icing conditions.
The heat transfer behavior of accreting ice surfaces in natural (flight test) and simulated (wind tunnel) cloud icing conditions have been studied. Observations of wet and dry ice growth regimes as measured by ultrasonic pulse-echo techniques were made. Observed wet and dry ice growth regimes at the stagnation point of a cylinder were compared with those predicted using a quasi steady-state heat balance model. A series of heat transfer coefficients were employed by the model to infer the local heat transfer behavior of the actual ice surfaces. The heat transfer in the stagnation region was generally inferred to be higher in wind tunnel icing tests than in natural, flight, icing conditions.
Flame configurations, flame-front cellular instability, and extinction of propane/air mixtures in the stagnation-point flow are experimentally studied for their dependence on downstream heat loss, preferential diffusion, and flame stretch. Boundaries for lean- and rich-limit extinction, stabilization of corrugated flames, and local extinction caused by sharp curvatures are mapped for varying propane concentrations and freestream velocities. Flame location and temperature at extinction are determined as functions of stagnation surface temperature, extent of preheating, propane concentration, and freestream velocity. Results substantiate the theoretical predictions of the different extinction modes for lean and rich flames in the absence of downstream heat loss, and yield useful insight on the extinction characteristics when finite downstream heat loss does exist. It is further shown that flame-front instability occurs only for rich mixtures in accordance with preferential diffusion considerations, and that flame stretch has a stabilizing effect such that flame-front instability is completely inhibited before the onset of extinction.
A material response model is developed for the determination of erosion rates and thermal response of woven materials to enable accurate thermal protection material sizing and future missions to Venus and giant planets. The stagnation-point thermal response is further evaluated using a Monte Carlo global sensitivity analysis to determine the interactions between key parameters in the material surface energy balance, such as, convective blowing correction parameter and weave material properties including isotropic and orthotropic thermal conductivites. The time-accurate material response for the dual layer weave erosion and surface temperatures show excellent agreement with stagnation flat face arc-jet tests involving time-varying aeroheating-cooling cycles. The high-fidelity Monte Carlo sensitivity analysis technique was used to address challenges in thermal protection sizing for the Adaptive Deployable Entry and Placement Technology (ADEPT) system, a low ballistic coefficient hypersonic decelerator. The results represent new correlations for such weaves involving Carbon-air oxidation equilibrium chemistry. Through the multi-variate regression analyses and uncertainty rankings performed, only three properties were found to contribute to the transient thermal response of the gore, with emissivity contributing to nearly 95% of the total output uncertainty.