High performance Solid Rocket Motor (SRM) submerged nozzle/combustion cavity flowfield assessment
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Two and three dimensional internal flowfield solutions for critical points in the Space Shuttle solid rocket booster burn time were developed using the Lockheed Huntsville GIM/PAID Navier-Stokes solvers. These perfect gas, viscous solutions for the high performance motor characterize the flow in the aft segment and nozzle of the booster. Two dimensional axisymmetric solutions were developed at t = 20 and t = 85 sec motor burn times. The t = 85 sec solution indicates that the aft segment forward inhibitor stub produces vortices with are shed and convected downwards. A three dimensional 3.5 deg gimbaled nozzle flowfield solution was developed for the aft segment and nozzle at t = 9 sec motor burn time. This perfect gas, viscous analysis, provided a steady state solution for the core region and the flow through the nozzle, but indicated that unsteady flow exists in the region under the nozzle nose and near the flexible boot and nozzle/case joint. The flow in the nozzle/case joint region is characterized by low magnitude pressure waves which travel in the circumferential direction. From the two and three dimensional flowfield calculations presented it can be concluded that there is no evidence from these results that steady state gas dynamics is the primary mechanism resulting in the nozzle pocketing erosion experienced on SRM nozzles 8A or 17B. The steady state flowfield results indicate pocketing erosion is not directly initiated by a steady state gas dynamics phenomenon.Keywords:
Solid-fuel rocket
Booster (rocketry)
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Isentropic process
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The performance of low-thrust rocket nozzles was studied with a full Navier-Stokes code. The effect of the reduction of the nozzle length on the viscous loss and on the two-dimensional loss due to the increase in the nozzle exit angle was examined by calculating the flowfield and performance values of hydrogen resistojet nozzle with various lengths and shapes (such as 20-deg or 30-deg conical nozzles and a nozzle whose wall contour is given by the Rao nozzle optimization code). It was found that the vacuum specific impulse value of the 30-deg conical nozzle was the highest and that of the contoured nozzle was the lowest among the three nozzles, whose throat Reynolds number and area ratio were 1150 and 82, respectively.
Rocket (weapon)
Rocket engine nozzle
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To reveal aft end internal flow feature of a solid rocket motor with submerged nozzle, a two dimension cold flow test model with rectangular channel was designed The cold gas simulation was based on geometric similarity and aerodynamic analogy Mean and fluctuating velocity field was measured using phase doppler particle analyzer (PDPA) It showed that the flow separates at upstream of the submerged nozzle and reattaches on the tip of the nozzle nose A relatively steady recirculation zone formed in the aft end cavity Both axial component and radial component of turbulence intensity were quite high
Solid-fuel rocket
Internal flow
Rocket (weapon)
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Abstract : An experimental program has been conducted in which unheated air was exhausted through a stationary rocket nozzle into a constant-area-tube launcher modified with a constrictive ring. The constrictive ring simulates a constraint which could be a permanent or temporary part of the launch tube. For the first phase of the test program the static wall-pressure distribution, upstream and downstream of the ring, and the mass flow-rates in the annular gap were measured. Given an underexpanded nozzle, a correlation between the generation blow-by, or reverse flow in the annular gap, and the position of the nozzle exit-plane relative to the front face of the ring was developed. The flow phenomena present in the launcher for different nozzle positions has also been investigated. For the second phase of the test program the static pressure distribution on the surface of the rocket and the mass flow-rates in the annular gap were measured. The differential pressure distributions on the rocket were related to the mass flow-rate in the annular gap with increasing blow-by flow, and with increasing ejector flow. (Author)
Rocket engine
Rocket (weapon)
Static pressure
Mass flow rate
Solid-fuel rocket
Mass flow
Rocket engine nozzle
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Gas generator
Mass flux
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The complex inner and outer flow-field of solid rocket motor(SRM) secondary hot gas injection vector nozzle is numerically investigated using three-dimensional average Reynolds N-S equations and the k-e turbulent models.The interaction between free outer flow and internal flow of vector nozzle is investigated.The results show that the flight parameters causes the change of pressure round the nozzle exit and affects the internal flow characteristics of the vector nozzle by means of the action of nozzle boundary layer.The side force has no connection with the flight status parameters when the flight Mach number is supersonic or the back pressure is lesser.On the low attitude and subsonic flow,the side force is reduced rapidly with the increasing of the flight Mach number and depressing of the back pressure.
Internal flow
Rocket engine nozzle
Discharge coefficient
External flow
Adverse pressure gradient
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A time-iterative full Navier-Stokes code, PARC, is used to analyze the flowfield of a two-dimensional ejector nozzle system. A parametric study was performed for two controlling parameters, duct to nozzle area ratio and nozzle pressure ratio. Results show that there is an optimum area ratio for the efficient pumping of secondary flow. At high area ratios, a freestream flow passes directly through the mixing duct without giving adequate pumping. At low area ratios, the jet boundary blocks the incoming flow. The nozzle pressure ratio variation shows that the pumping rate increases as the pressure ratio increases, provided there is no interaction between the shroud wall and the shock cell structure.
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A submerged annular conical aerospike nozzle for a design Mach number of 2.0 has been numerically studied under over-expanded condition. Particular attention has been paid to the flow separation characteristics in aerospike nozzle for underwater propulsion application. The detailed Navier- Stokes flow computations were utilized to elucidate the gas-water interactions under the framework of a Volume of fluid model. The calculation results show that,the intermittently necking/bulging or necking/back-attack phenomenon also exists in the submerged over-expanded annular conical aerospike nozzle flowfield. The back pressure for underwater nozzle is not only determined by the pressure of the water environment but also by the pressure in the gas bubble,and the gas/water interface restricts the gaseous jet as a wall, which is extremely different from the air environment condition. The gas/water interface around the plug flowfield is observed to be severely affected by the pulsation of the nozzle expansion back pressure,and shows continual oscillation between necking and bulging. The flow separation characteristics develop and change along with the jet oscillation. Depending on shock structures in the nozzle and flow separation characteristic on the plug surface,the nozzle exhibits different flow separation regimes which can be broadly classified into five types. Further,the mechanism for unsteady separation in submerged over-expanded aerospike nozzle is different from the one under air environment condition. The predicted wall fluctuating pressure shows a broad-frequency phenomenon with the bandwidth of 0~1000Hz by the current Navier-Stokes computation. As a result of the complicated coupling effects among the back pressure pulsation,the flow separation process and the jet oscillation,flow separation behavior in the present annular conical aerospike nozzle for underwater propulsion shows highly irregular oscillation characteristics.
Back pressure
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A time-dependent technique, in conjunction with the boundary-fitted coordinates system, is applied to solve a gas-only one-phase flow and a fully-coupled, gas-particle two-phase flow inside nozzles with small throat radii of curvature, steep wall gradients, and submerged configurations. The emphasis of the study has been placed on one- and two-phase flow in the transonic region. Various particle sizes and particle mass fractions have been investigated in the two-phase flow. The salient features associated with the two-phase nozzle flow compared with those of the one-phase flow are illustrated through the calculations of the JPL nozzle, the Titan III solid rocket motor, and the submerged nozzle configuration found in the Inertial Upper Stage (IUS) solid rocket motor.
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Discharge coefficient
Rocket engine nozzle
Overall pressure ratio
Rocket (weapon)
Rocket engine
Adverse pressure gradient
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