ALBERT

Description: A review of the relevant flight conditions and physical models for planetary probe afterbody aeroheating calculations is given. Readily available sources of afterbody flight data and published attempts to computationally simulate those flights are summarized. A current status of the application of turbulence models to afterbody flows is presented. Finally, recommendations for additional analysis and testing that would reduce our uncertainties in our ability to accurately predict base heating levels are given.

Description: A three-dimensional laser Doppler velocimeter has laser optics for a first channel positioned to create a probe volume in space, and laser optics and for second and third channels, respectively, positioned to create entirely overlapping probe volumes in space. The probe volumes and overlap partially in space. The photodetector is positioned to receive light scattered by a particle present in the probe volume, while photodetectors and are positioned to receive light scattered by a particle present in the probe volume. The photodetector for the first channel is directly connected to provide a first channel analog signal to frequency measuring circuits. The first channel is therefore a primary channel for the system. Photodetectors and are respectively connected through a second channel analog signal attenuator to frequency measuring circuits and through a third channel analog signal attenuator to frequency measuring circuits. The second and third channels are secondary channels, with the second and third channels analog signal attenuators and controlled by the first channel measurement burst signal on line. The second and third channels analog signal attenuators and attenuate the second and third channels analog signals only when the measurement burst signal is false.

Description: A laser Doppler velocimeter multiplexer interface includes an event pulse synchronizer which synchronizes data pulses from events A, B, and C. Clock control is connected to receive timing information on the data pulses from the synchronizer. Displays are connected to receive clock signals from the clock control for indicating a data rate for each of the measured events A, B, and C. The display is connected to receive clock signals from the clock control to indicate a coincidence rate between data pulses for any selected combination of the measured events A, B, and C. A multiplexer receives the data pulses from the events A, B, and C and rate data from the clock control. The multiplexer has output for supplying the data pulses and rate data to a single input of a data processing system. A multiplexer control is connected to supply control signals to the multiplexer for selecting the event data pulses and the rate data for output from the multiplexer. The multiplexer control receives start signals from the pulse synchronizer and user selected inputs for desired outputs from the multiplexer.

Description: An aerodynamic holographic interferometry system has been developed and adapted for use with an existing supersonic wind tunnel. The basis for the system is briefly discussed, and the flow investigation for which the system was developed is outlined. The system is described, including detailed equipment schematics and cost figures. Some details and guidance are given with regard to experience in reducing the interferometry fringe data via the Abel transformation to obtain density fields and, in turn, velocity fields. Finally, representative results are shown together with complementary laser velocimetry measurements and Navier-Stokes computations.

Description: An experimental study of the axisymmetric shock-wave / boundary-layer strong interaction flow generated in the vicinity of a cylinder-cone intersection was conducted. The study data are useful in the documentation and understanding of compressible turbulent strong interaction flows, and are part of a more general effort to improve turbulence modeling for compressible two- and three-dimensional strong viscous/inviscid interactions. The nominal free stream Mach number was 2.85. Tunnel total pressures of 1.7 and 3.4 atm provided Reynolds number values of 18 x 10(6) and 36 x 10(6) based on model length. Three cone angles were studied giving negligible, incipient, and large scale flow separation. The initial cylinder boundary layer upstream of the interaction had a thickness of 1.0 cm. The subsonic layer of the cylinder boundary layer was quite thin, and in all cases, the shock wave penetrated a significant portion of the boundary layer. Owing to the thickness of the cylinder boundary layer, considerable structural detail was resolved for the three shock-wave / boundary-layer interaction cases considered. The primary emphasis was on the application of the holographic interferometry technique. The density field was deduced from an interferometric analysis based on the Able transform. Supporting data were obtained using a 2-D laser velocimeter, as well as mean wall pressure and oil flow measurements. The attached flow case was observed to be steady, while the separated cases exhibited shock unsteadiness. Comparisons with Navier-Stokes computations using a two-equation turbulence model are presented.

Description: The present paper provides the background of a focused effort to assess uncertainties in predictions of heat flux and pressure in hypersonic flight (airbreathing or atmospheric entry) using state-of-the-art aerothermodynamics codes. The assessment is performed for four mission relevant problems: (1) shock turbulent boundary layer interaction on a compression corner, (2) shock turbulent boundary layer interaction due a impinging shock, (3) high-mass Mars entry and aerocapture, and (4) high speed return to Earth. A validation based uncertainty assessment approach with reliance on subject matter expertise is used. A code verification exercise with code-to-code comparisons and comparisons against well established correlations is also included in this effort. A thorough review of the literature in search of validation experiments is performed, which identified a scarcity of ground based validation experiments at hypersonic conditions. In particular, a shortage of useable experimental data at flight like enthalpies and Reynolds numbers is found. The uncertainty was quantified using metrics that measured discrepancy between model predictions and experimental data. The discrepancy data is statistically analyzed and investigated for physics based trends in order to define a meaningful quantified uncertainty. The detailed uncertainty assessment of each mission relevant problem is found in the four companion papers.

Description: The surface interferometric skin-friction (SISF) measurement system is an instrument for determining the distribution of surface shear stress (skin friction) on a wind-tunnel model. The SISF system utilizes the established oil-film interference method, along with advanced image-data-processing techniques and mathematical models that express the relationship between interferograms and skin friction, to determine the distribution of skin friction over an observed region of the surface of a model during a single wind-tunnel test. In the oil-film interference method, a wind-tunnel model is coated with a thin film of oil of known viscosity and is illuminated with quasi-monochromatic, collimated light, typically from a mercury lamp. The light reflected from the outer surface of the oil film interferes with the light reflected from the oil-covered surface of the model. In the present version of the oil-film interference method, a camera captures an image of the illuminated model and the image in the camera is modulated by the interference pattern. The interference pattern depends on the oil-thickness distribution on the observed surface, and this distribution can be extracted through analysis of the image acquired by the camera. The oil-film technique is augmented by a tracer technique for observing the streamline pattern. To make the streamlines visible, small dots of fluorescentchalk/oil mixture are placed on the model just before a test. During the test, the chalk particles are embedded in the oil flow and produce chalk streaks that mark the streamlines. The instantaneous rate of thinning of the oil film at a given position on the surface of the model can be expressed as a function of the instantaneous thickness, the skin-friction distribution on the surface, and the streamline pattern on the surface; the functional relationship is expressed by a mathematical model that is nonlinear in the oil-film thickness and is known simply as the thin-oil-film equation. From the image data acquired as described, the time-dependent oil-thickness distribution and streamline pattern are extracted and by inversion of the thin-oil-film equation it is then possible to determine the skin-friction distribution. In addition to a quasi-monochromatic light source, the SISF system includes a beam splitter and two video cameras equipped with filters for observing the same area on a model in different wavelength ranges, plus a frame grabber and a computer for digitizing the video images and processing the image data. One video camera acquires the interference pattern in a narrow wavelength range of the quasi-monochromatic source. The other video camera acquires the streamline image of fluorescence from the chalk in a nearby but wider wavelength range. The interference- pattern and fluorescence images are digitized, and the resulting data are processed by an algorithm that inverts the thin-oil-film equation to find the skin-friction distribution.

Description: Over the past 20 years, the use of oil-film interferometry to measure the skin friction coefficient (C(sub f) = tau/q where tau is the surface shear stress and q is the dynamic pressure) has increased. Different forms of this oil-film technique with various levels of accuracy and ease of use have been successfully applied in a wide range of flows. The method's popularity is growing due to its relative ease of implementation and minimal intrusiveness as well as an increased demand for C(sub f) measurements. Nonetheless, the accuracy of these methods has not been rigorously addressed to date. Most researchers have simply shown that the skin-friction measurements made using these techniques compare favorably with other measurements and theory, most of which are only accurate to within 5-20%. The use of skin-friction data in the design of commercial aircraft, whose drag at cruise is 50% skin-friction drag, and in the validation of computational fluid dynamics programs warrants better uncertainty estimates. Additional information is contained in the original extended abstract.