Static performance tests of a flight-type STOVL ejector

The design and development of thrust augmenting STOVL ejectors has typically been based on experimental iteration (i.e., trial and error). Static performance tests of a full scale vertical lift ejector were performed at primary flow temperatures up to 1560 R (1100 F). Flow visualization (smoke generators and yarn tufts) were used to view the inlet air flow, especially around the primary nozzle and end plates. Performance calculations are presented for ambient temperatures close to 480 R (20 F) and 535 R (75 F) which simulate seasonal aircraft operating conditions. Resulting thrust augmentation ratios are presented as functions of nozzle pressure ratio and temperature

A full-scale STOVL ejector experiment

The design and development of thrust augmenting short take-off and vertical landing (STOVL) ejectors has typically been an iterative process. In this investigation, static performance tests of a full-scale vertical lift ejector were performed at primary flow temperatures up to 1560 R (1100 F). Flow visualization (smoke generators, yarn tufts and paint dots) was used to assess inlet flowfield characteristics, especially around the primary nozzle and end plates. Performance calculations are presented for ambient temperatures close to 480 R (20 F) and 535 R (75 F) which simulate 'seasonal' aircraft operating conditions. Resulting thrust augmentation ratios are presented as functions of nozzle pressure ratio and temperature. Full-scale experimental tests such as this are expensive, and difficult to implement at engine exhaust temperatures. For this reason the utility of using similarity principles -- in particular, the Munk and Prim similarity principle for isentropic flow -- was explored. At different primary temperatures, exit pressure contours are compared for similarity. A nondimensional flow parameter is then shown to eliminate primary nozzle temperature dependence and verify similarity between the hot and cold flow experiments. Under the assumption that an appropriate similarity principle can be established, then properly chosen performance parameters should be similar for both hot flow and cold flow model tests

Use of an approximate similarity principle for the thermal scaling of a full-scale thrust augmenting ejector

Full temperature ejector model simulations are expensive, and difficult to implement experimentally. If an approximate similarity principle could be established, properly chosen performance parameters should be similar for both hot and cold flow tests if the initial Mach number and total pressures of the flow field are held constant. Existing ejector data is used to explore the utility of one particular similarity principle; the Munk and Prim similarity principle for isentropic flows. Static performance test data for a full-scale thrust augmenting ejector are analyzed for primary flow temperatures up to 1560 R. At different primary temperatures, exit pressure contours are compared for similarity. A nondimensional flow parameter is then used to eliminate primary nozzle temperature dependence and verify similarity between the hot and cold flow experiments

Evaluation of F/A-18A HARV inlet flow analysis with flight data

The F/A-18A aircraft has experienced engine stalls at high angles-of-attack and yaw flight conditions which were outside of its flight envelope. Future aircraft may be designed to operate routinely in this flight regime. Therefore, it is essential that an understanding of the inlet flow field at these flight conditions be obtained. Due to the complex interactions of the fuselage and inlet flow fields, a study of the flow within the inlet must also include external effects. Full Navier-Stokes (FNS) calculations on the F/A-18A High Alpha Research Vehicle (HARV) inlet for several angles-of-attack with sideslip and free stream Mach numbers have been obtained. The predicted forebody/fuselage surface static pressures agreed well with flight data. The surface static pressures along the inlet lip are in good agreement with the numerical predictions. The major departure in agreement is along the bottom of the lip at 30 deg and 60 deg angle-of-attack where a possible streamwise flow separation is not being predicted by the code. The circumferential pressure distributions at the engine face are in very good agreement with the numerical results. The variation in surface static pressure in the circumferential direction is very small with the exception of 60 angle-of-attack. Although the simulation does not include the effect of the engine, it appears that this omission has a second order effect on the circumferential pressure distribution. An examination of the unsteady flight test data base has shown that the secondary vortex migrates a significant distance with time. In fact, the extent of this migration increases with angle-of-attack with increasing levels of distortion. The effects of the engine on this vortex movement is unknown. This implies that the level of flow unsteadiness increases with increasing distortion. Since the computational results represent an asymptotic solution driven by steady boundary conditions, these numerical results may represent an arbitrary point in time. A comparison of the predicted total pressure contours with flight data indicates that the numerical results are within the excursion range of the unsteady data which is the best the calculations can attain unless an unsteady simulation is performed