Wind Tunnel Testing of a One-Dimensional Laser Beam Scanning and Laser Sheet Approach to Shock Sensing

Abstract A 15- by 15-cm supersonic wind tunnel application of a one-dimensional laser beam scanning approach to shock sensing is presented. The measurement system design allowed easy switching between a focused beam and a laser sheet mode for comparison purposes. The scanning results were compared to images from the tunnel Schlieren imaging system. The tests revealed detectable changes in the laser beam in the presence of shocks. The results lend support to the use of the one-dimensional scanning beam approach for detecting and loc ating shocks in a flow, but some issues must be addressed in regards to noise and other limitations o f the system. Introduction Airborne vehicles moving through the atmosphere with a speed above that of sound may generate shock waves. The shock waves may be either detached from the flying vehicle like, for instance, bow shocks or attached to the vehicle’s internal or external surfaces like normal or oblique shocks. Formation and propagation of aerodynamic shocks are important considerations in the design of in-flight control systems and in the performance evaluation of individual components in ground test facilities (Refs. 1 to 3). Moreover, monitoring of aerodynamic shocks and their interactions enhances the safety and performance of supersonic and hypersonic flights (Refs. 4 and 5). The need to develop shock position sensors capable of meeting flight qualifying requirements was especially emphasized during the High Speed Commercial Transport (HSCT) development (Refs. 6 and 7). In the process, various approaches were analyzed and sensor prototypes were developed and tested (Refs. 8 and 9). Early efforts were concentrated around using pressure taps along the inlet walls. The positions of the shocks were determined by tracking the pressure reading and locating the pressure jump associated with the shock. This basic technique evolved in several wall pressure-based configurations of normal shock position sensing systems. Despite an apparent initial success, these wall pressure-based measuring techniques had serious drawbacks. The most important drawbacks were slow response due to pneumatic manifolds used and the effect of the boundary layer on the stability of pressure readings. These problems seriously restricted applicability of these techniques to normal shock detection and control during supersonic flight. Moreover, for a commercial aircraft, economic efficiency has to be achieved in order to make supersonic flight economically viable. As a result, the flight control system is required, in addition to avoiding a un-start, to provide the most economical operating regime for the engine (achieved by minimizing the fuel consumption). Optical flow visualization has been widely used in ground-based flow analyzing facilities. Instrumentation and measurement schemes used in these cases have typically included well-known flow visualization techniques such as shadowgraph, Schlieren, and interferometry (Refs. 10 to 13). Advantages