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The objective of this research is to understand supersonic laminar flow stability, transition and active control. Some prediction techniques are developed or modified to analyze laminar flow stability. The effects of distributed heating and cooling as an active boundary layer control technique are studied. The primary tasks of the research apply to the NASA/Ames Proof-of-Concept (PoC) and the Laminar Flow Supersonic Wind Tunnel's (LFSWT's) nozzle design with laminar flow control and are listed as follows: (1) Predictions of supersonic laminar boundary layer stability and transition; (2) Effects of wall heating and cooling on supersonic laminar flow control on a flat plate; (3) Performance evaluation of the PoC and LFSWT nozzle designs with wall heating and cooling applied at different locations and various lengths; (4) Effects of a conducted -vs- pulse wall temperature distribution for the LFSWT; and (5) Application of wall heating and/or cooling to laminar boundary layer and flow separation control of airfoils and investigation of related active control techniques.

A Wall Signature method originally developed by Hackett has been selected to be adapted for the Ames 12-ft Wind Tunnel WIAC system in the project. This method uses limited measurements of the static pressure at the wall, in conjunction with the solid wall boundary condition, to determine the strength and distribution of singularities representing the test article. The singularities are used in turn for estimating wall interference at the model location. The development and implementation of a working prototype will be completed, delivered and documented with a software manual. The WIAC code will be validated by conducting numerically simulated experiments rather than actual wind tunnel experiments. The simulations will be used to generate both free-air and confined wind-tunnel flow fields for each of the test articles over a range of test configurations. Specifically, the pressure signature at the test section wall will be computed for the tunnel case to provide the simulated 'measured' data. These data will serve as the input for the WIAC method--Wall Signature method. The performance of the WIAC method then may be evaluated by comparing the corrected parameters with those for the free-air simulation. The following two additional tasks are included: (1) On-line wall interference calculation: The developed wall signature method (modified Hackett's method) for Ames 12-ft Tunnel will be the pre-computed coefficients which facilitate the on-line calculation of wall interference, and (2) Support system effects estimation: The effects on the wall pressure measurements due to the presence of the model support systems will be evaluated.

A wall signature method has been selected to be adapted for the Ames 12-ft Wind Tunnel WIAC system. This uses limited measurements of the static pressure at the wall, in conjunction with the solid wall boundary condition, to determine the strength and distribution of singularities representing the test article. The singularities are used to estimate wall interference at the model location. The development and implementation of a working prototype will be completed, delivered and documented with a software manual.

The objective of the research is to understand supersonic laminar flow stability, transition and active control. Some prediction techniques will be developed or modified to analyze laminar stability. The effects of supersonic laminar flow with distributed heating and cooling on active control will be studied. The primary tasks of the research applying to the NASA/Ames POC and LFSWT's nozzle design with laminar flow control are as follows: (1) supersonic laminar boundary layer stability and transition prediction; (2) effects of heating and cooling for supersonic laminar flow control; and (3) POC and LFSWT nozzle design with heating and cooling effects combining wall contour and length changes.

A prototype software system was constructed to detect anomalous Space Shuttle Main Engine (SSME) behavior in the early stages of fault development significantly earlier than the indication provided by either redline detection mechanism or human expert analysis. The major task of the research project is to analyze ground test data, to identify patterns associated with the anomalous engine behavior, and to develop a pattern identification and detection system on the basis of this analysis. A prototype expert system which was developed on both PC and Symbolics 3670 lisp machine for detecting anomalies in turbopump vibration data was checked with data from ground tests 902-473, 902-501, 902-519, and 904-097 of the Space Shuttle Main Engine. The neural networks method was also applied to supplement the statistical method utilized in the prototype system to investigate the feasibility in detecting anomalies in turbopump vibration of SSME. In most cases the anomalies detected by the expert system agree with those reported by NASA. On the neural networks approach, the results are given the successful detection rate higher than 95 percent to identify either normal or abnormal running condition based on the experimental data as well as numerical simulation.

The statistical and neural networks methods have been applied to investigate the feasibility in detecting anomalies in turbopump vibration of SSME. The anomalies are detected based on the amplitude of peaks of fundamental and harmonic frequencies in the power spectral density. These data are reduced to the proper format from sensor data measured by strain gauges and accelerometers. Both methods are feasible to detect the vibration anomalies. The statistical method requires sufficient data points to establish a reasonable statistical distribution data bank. This method is applicable for on-line operation. The neural networks method also needs to have enough data basis to train the neural networks. The testing procedure can be utilized at any time so long as the characteristics of components remain unchanged.

An improved wall interference assessment and correction method for three-dimensional wind tunnel testing is presented. Blockage corrections on the surface of a test article are calculated based on a limited number of wall pressure measurements. These measurements are combined with a signature analysis procedure and influence functions to determine an equivalent test article and wake representation. Pressure coefficient corrections are calculated based on this equivalent body. The signature analysis procedure is modified to improve the on-line operation of the wall signature method. A new geometry of the equivalent body is introduced which can be combined with existing panel codes more effectively. The calculation of influence functions and the determination of pressure coefficient corrections are based on a panel code. Therefore it is possible to apply the present method to any closed tunnel cross section. A numerical simulation of the idealized flow field of a wing and its wake in a rectangular wind tunnel is used to verify the improved wall signature method. The present method is considered to calculate blockage corrections in the NASA/ARC 12ft Pressure Wind Tunnel.