For more than a half-century, several types of altitude-compensating rocket nozzles have been proposed and analyzed, but very few have been adequately tested in a relevant flight environment. One type of altitude-compensating nozzle is the dual-bell rocket nozzle, which was first introduced into literature in 1949. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. This paper proposes a method for conducting testing and research with a dual-bell rocket nozzle in a flight environment. We propose to leverage the existing NASA F-15 airplane and Propulsion Flight Test Fixture as the flight testbed, with the dual-bell nozzle operating during captive-carried flights, and with the nozzle subjected to a local flow field similar to that of a launch vehicle. The primary objective of this effort is not only to advance the technology readiness level of the dual-bell nozzle, but also to gain a greater understanding of the nozzle mode transitional sensitivity to local flow-field effects, and to quantify the performance benefits with this technology. The predicted performance benefits are significant, and may result in reducing the cost of delivering payloads to low-Earth orbit. Nomenclature
An axisymmetric full Navier-Stokes computational fluid dynamics study is conducted to examine nozzle exhaust jet plume effects on the sonic boom signature of a supersonic aircraft. A simplified axisymmetric nozzle geometry, representative of the nozzle on the NASA Dryden NF-15B Lift and Nozzle Change Effects on Tail Shock research airplane, is considered. The computational fluid dynamics code is validated using available wind-tunnel sonic boom experimental data. The effects of grid size, spatial order of accuracy, grid type, and flow viscosity on the accuracy of the predicted sonic boom pressure signature are quantified. Grid lines parallel to the Mach wave direction are found to give the best results. Second-order accurate upwind methods are required as a minimum for accurate sonic boom simulations. The highly underexpanded nozzle flow is found to provide significantly more reduction in the tail shock strength in the sonic boom N-wave pressure signature than perfectly expanded and overexpanded nozzle flows. A tail shock train in the sonic boom signature is observed for the highly underexpanded nozzle flow. Axisymmetric computational fluid dynamics simulations show the flow physics inside the F-15 nozzle to be nonisentropic and complex. Although the one-dimensional isentropic nozzle plume results look reasonable, they fail to capture the sonic boom shock train in the highly underexpanded nozzle flow. Nomenclature
Flight research has been conducted on an aerospike rocket nozzle using high power solid rockets. Two aerospike rockets and one conventional rocket were flown successfully to supersonic speeds, providing the first known set of transonic flight performance data for aerospike rockets. This paper describes the rockets, solid rocket motors, nozzles, and rocket instrumentation system. Flight test results are also discussed and compared with ground test results. Flight data show that all of the rockets successfully reached supersonic speeds with a maximum Mach number of 1.6 and a peak pressure altitude of nearly 30,000 ft. The aerospike nozzle efficiency was determined to be 0.96 from computational fluid dynamics (CFD) analysis. The rocket chamber pressures and thrusts of the aerospike rocket motors were lower than the conventional rocket motors. Because the same propellant formulation was used in all of the rocket motors, the discrepancy in pressure and thrust was most likely caused by a larger actual aerospike nozzle throat area than the designed throat area. Potential causes for the larger aerospike nozzle throat area are also discussed.nozzle coefficient of thrust CFD = computational fluid dynamics 2 d = diameter D = external aerodynamic drag of the rocket FADS = flush air data system F x = forces in the axial direction g = gravitational acceleration ISP = specific impulse M = Mach number m = rocket mass m & = rocket propellant mass flow rate P a = ambient pressure P c = rocket chamber pressure psig = pounds per square inch gage r = radius T = rocket motor thrust t = time u x = rocket velocity in the axial direction U ∞ = freestream velocity x = axial direction γ= specific heat ratio η t = nozzle efficiency factor ρ ∞ = freestream density θ = rocket elevation angle I. Introduction HE broad range of human and robotic space exploration missions to the Moon, Mars, and destinations beyond, as outlined in the President's Space Exploration Vision, will benefit from advances in technology from all systems of a space vehicle, including the propulsion system and its nozzle. Most space vehicles, including spacelaunch boosters, Moon and Mars landers, orbiters, space tugs, or deep-space interplanetary probes, rely on a nozzle to convert the energy generated in the propulsion system into the direct thrust necessary to propel them through space. This is true for chemical, nuclear thermal, solar-heating, and arc-heating rockets. Therefore, nozzle efficiency and its physical size have a primary effect on the overall performance of space vehicles. Increased nozzle efficiencies and shorter nozzle lengths translate directly into less weight, more payload, longer range, and lower costs across the entire spectrum of space vehicles.Reference 1 provides a broad review of advanced rocket nozzle technologies. In this reference, it is shown that altitude-compensating nozzle technologies, such as the expansion-deflection nozzle, aerospike or plug nozzle, dualbell nozzle, and dual-expander nozzle, can provide significant performance gains by a...
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