Low Nozzle Pressure Ratios Research Articles

Lobed Mixers Using Simultaneous Laser-Induced Fluorescence and Mie Scattering

Fig. 5 Thrust coefŽ cient predictions. experimental results in predicting the pressure variation behind the shock. In general, the two-equation turbulence models gave the best overall agreement with the experimental results for the pressure distribution and thrust coefŽ cient at overexpanded conditions. Figure 4 compares the computed pressure distribution over the  ap using the k-v model to the experimental results at the centerline and near the end wall at Ž ve overexpanded NPRs. The computed thrust coefŽ cients using the two-equation turbulence models are compared to the experimental results in Fig. 5 and Table 1. One can see that the thrust coefŽ cient was predicted within 1.0 and 2% of the experimental values for NPRs above 50 and 30% design, respectively. At lower nozzle pressure ratios, two-dimensional  ow predictions are inadequate because of the strong three-dimensional  ow effects behind the shock, which are observed in the experimental results.

Comparison of computational and experimental jet effects

Computations were made for transonic flow over a nozzle afterbody with a real plume. The effects of parameters such as nozzle pressure ratio, plume stagnation temperature, and interior nozzle shape on afterbody drag were quantified and compared with available experimental data. Remarkable agreement between the computational and experimental results is demonstrated, including the drag minimum associated with low nozzle pressure ratios and the absolute dependence of drag on temperature. The computer program used for the calculations is an axisymmetric modification of Deiwert's planar Navier-Stokes program and utilizes a mixed explicit-implicit MacCormack algorithm with an algebraic turbulence model.

On the Structure of Jets from Highly Underexpanded Nozzles Into Still Air

A method for calculating the position of the first normal shock, or Mach disc, in the jet behind a highly underexpanded nozzle is presented. In the calculation for a sonic orifice, the axial pressure distribution on the centerline of the flow behind the orifice, calculated by the method of characteristics, is used to define a fictitious nozzle extension, and the shock is then assumed to exist at that point where atmospheric pressure would be attained behind the shock—i.e., the shock is assumed to exist at the end of the fictitious nozzle extension. Physical arguments are then employed to extend this calculation to nozzles with supersonic exit Mach Numbers. The results compare favorably with experimental data. An approximate method for computing the jet boundary up to the point of maximum jet area is given, and the results are compared both with photographs of actual jets and with jet boundaries calculated by the method of characteristics. Favorable agreement exists at relatively low nozzle pressure ratios.