unpublished) and fluidic throat-area control of axisymmetric nozzles has been investigated (Catt et al.), the axisymmetric nozzle has yet to be adequately explored for its potential in fluidic thrust-vector control for air-breathing propulsion systems. A fluidic thrust-vectoring nozzle with axisymmetric geometry has the benefits of structural efficiency and the possibility of retrofit onto existing aircraft.

A sub-scale experimental static investigation of an axisymmetric nozzle with fluidic injection for thrust vectoring was conducted at the NASA Langley Jet Exit Test Facility. Fluidic injection was introduced through flush-mounted injection ports in the divergent section. Variations on the injection port design included slot longitudinal location, number of slots, and a distributed pattern of holes rather than slots. Test conditions included a range of nozzle pressure ratios (NPR) from 2 to 10 and a range of injection total pressure ratio (SPR) from no-flow to 1.5. Measurements included normal and axial forces and internal static pressures distributed over a 180? arcsector of the nozzle (nearly full internal coverage due to symmetry) extending from upstream of the nozzle throat to the nozzle trailing edge. A paint flow visualization technique was applied for selected configurations and test conditions: an oilbased paint was applied to the divergent-duct internal surface and allowed to dry at constant nozzle operating conditions. The dried paint patterns indicate flow features at the nozzle surface.

Model Description

The model was a 10 percent-scale axisymmetric nozzle with a throat area of 3.736 in2 and an expansion ratio (exit area divided by throat area) of 1.738. Two independently-operated injection ports were located in opposing sectors of the divergent duct for fluidic thrust vectoring. For the data presented in this paper, only one port was flowing; the other port existed only as a closed cavity. An additional aspect of this experiment (not reported on here) was the investigation of symmetric injection for expansion control, where both slots were flowing.

A sketch of the model hardware is shown in figure 1. The main hardware components consisted of a convergent section, a throat-block insert, an injection plenum disk, several pairs of interchangeable injection-port inserts, and a plenum cover. The injection plenum disk contained two isolated 90?-sector opposing plenums for fluidic injection. Each plenum contained a baffle plate for even pressure distribution and three static pressure taps for plenum pressure measurement. The plenums were designed to accept interchangeable inserts for variations in injection-port geometry. The injection-port geometry variations included slot axial location (forward or aft), numbers of slots (single or triple), and a distributed hole pattern rather than slots. All of the injection ports were tested with injection flow delivered across an arc angle of 60? (?30? off q=0?).

The nozzle internal surface was instrumented with 57 static pressure orifices arranged in seven longitudinal rows spaced 30?

apart, with q=0? centered on the upper (flowing) injection slot. The orifice rows extended from just upstream of the nozzle throat to near the trailing edge, resulting in nearly complete coverage of the internal divergent duct due to nozzle symmetry.

Facility Description

The test was conducted in the NASA Langley Jet Exit Test Facility (JETF). The facility is used for static performance testing of scaled nozzle concepts and is capable of supplying multiple flows to the nozzle. The JETF consists of two independently supplied and controlled high-pressure air lines, a dual-flow propulsion simulation stand (shown in figure 2), and a data acquisition system. Each air supply can be continuously regulated up to approximately 23 lbf/sec. The air supply is maintained at approximately 535?R during testing, and the flow is exhausted through the nozzle to atmospheric pressure in a large, vented room. Further information on the JETF can be found in NASA-TM-102750 (see References).

Results and Discussion

Overexpanded Primary Jet

The system resultant thrust coefficient (Cf,gsys), based on

the sum of the ideal thrusts of the primary and injected flow, and the thrust-vector angle (dv) of the aft-slot configuration as

functions of secondary weight flow ratio (w?t) for NPR=3 and NPR=8.26 are presented in figure 3. The lower NPR represents a low-speed operating condition such as takeoff or high angleof-attack flight where thrust vectoring might be used to supplement or replace the conventional aerodynamic controls. The higher NPR is the design pressure ratio for this configuration (based on the exit-to-throat-area expansion ratio) where thrust vectoring might be used for trim-drag reduction at cruise operating conditions.

At NPR=3, the nozzle was producing thrust at a relatively low performance level (Cf,gsys=0.895), expected behavior of a

nozzle operating at significantly overexpanded flow conditions. In an operational nozzle, additional steps would be taken to ensure that the nozzle thrust efficiency would be raised to higher levels (such as the fluidic expansion control technique also investigated in this experiment, but not reported on in this paper). At the lower NPR, thrust-vector angles of up to dv=18?

were generated with mainly positive impact on system resultant thrust coefficient. For w?t ? 0.08, an increase in performance over that for no injection was measured.

The pressure distribution at NPR=3 is shown in figure 4; the solid symbols represent the ?clean? nozzle configuration with no injection slots, and the open symbols represent the aftslot configuration at a secondary pressure ratio (SPR) of 1. The