Modeling of Two-Stage Ejector for High-Altitude Testing of Satellite Thrusters
28 Aug 2012-AIAA Journal (American Institute of Aeronautics and Astronautics (AIAA))-Vol. 50, Iss: 6, pp 1398-1408
TL;DR: In this paper, the performance of a two-stage external ejector during high-altitude testing of large-area-ratio satellite thrusters is numerically investigated, and the predicted results compare well with in-house experimental data.
Abstract: The performance of a two-stage ejector during high-altitude testing of large-area-ratio satellite thrusters is numerically investigated. Since theflowrate of the exhaust from the satellite thruster is very low, self-ejector action of the exhaust is quite weak; therefore, a two-stage external ejector is required to create the desired low vacuum in the test chamber. The present work attempts to investigate the effects of various operational and geometric parameters on the performance of the two-stage ejector. Predicted results show that the downstream ejector (E2) operation is more critical for maintaining the required vacuum. However, for optimal performance, it is possible to tune the parameters such that both ejectors deliver the same suction effect.Maximumperformance of each ejector is obtained when the primary jet expanding from the nozzle smoothly seals the mixer throat against backflow. Employing a low molecular weight fluid and high stagnation temperature helps in reducing the quantity of fluid required for test facility evacuation. Useful correlations have been derived to quantify the suction performance of the two-stage ejector, and the predicted results compare well with in-house experimental data.
Citations
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TL;DR: In this article, the authors evaluated the effect of an annular rocket exhaust pattern on the mixing characteristics of a simple ejector and found that it yields a ratio of entrained air to rocket mass flows over 75% higher than an equivalent circular rocket exhaust nozzles over a range of rocket chamber total pressures.
Abstract: Experimental tests are conducted to evaluate the effect of an annular rocket exhaust pattern on the mixing characteristics of a simple ejector. Rocket exhaust is simulated using pure oxygen while air is entrained from the surroundings at static conditions. Results are compared with an equivalent configuration using traditional circular rocket exhaust nozzles. It is shown that an annular rocket exhaust pattern yields a ratio of entrained air to rocket mass flows over 75% higher than an equivalent circular rocket exhaust pattern over a range of rocket chamber total pressures. It is demonstrated that an annular rocket exhaust pattern within a straight ejector duct with a length-to-diameter ratio of 6 is able to produce entrained air to rocket mass flow ratios up to 10% higher than a straight then expanding ejector configuration twice as long employing a circular rocket exhaust pattern. Exit plane measurements of both total pressure and oxygen concentration are also collected at high and low rocket chamber to...
12 citations
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...[12]) and rocket engines (German et al....
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TL;DR: In this article, the mixing of two supersonic streams in a SU-personic exhaust diffuser using theoretical, numerical, and experimental approaches was investigated, and the effects of diffuser contraction ratio (CR) and convergence angle (θ) on the starting process, pressure recovery, and flow features were investigated.
Abstract: This paper investigates the mixing of two supersonic streams in a supersonic exhaust diffuser using theoretical, numerical, and experimental approaches. It has focused attention on the startup and stationary operation of the two-stream supersonic diffusers. The diffuser geometry consists of a constant area throat which connects two ducts of convergent and divergent cross-sections. The effects of diffuser contraction ratio (CR) and convergence angle (θ) on the starting process, pressure recovery, and flow features. Experiments have been performed at five different secondary Mach number ( M s ) values in the range of 1.8 M s M p ) value of 2. Synchronized pressure measurements are used to study the diffuser performance characteristics. Numerical simulations were also carried out to obtain a better insight into the flow field and computed results have been compared with the experimental results. It was found that the starting behavior has a hysteresis. As a result, the tunnel can be operated at lower pressure ratio than the starting pressure ratio. It is observed that the settling chamber pressure required to start a Constant-Area Diffuser (CAD) is decreased with increase in the diffuser length up to an L / D ratio of 9; subsequently, starting pressure remains almost constant with the further increase in the length due to frictional and pressure recovery losses. The present study provides valuable insights into the multi-stream supersonic mixing and diffusion problem.
8 citations
TL;DR: In this paper, the performance of a cold gas thruster operating under the vacuum with a high expansion ratio was investigated, and the role of second throat type diffuser in the overall pressure recovery was also investigated.
Abstract: The present study is carried out to investigate the performance of a cold gas thruster operating under the vacuum with a high expansion ratio. A pressure recovery system employing diffuser and multistage ejectors is employed to evacuate a vacuum cell whose pressure is an indicating parameter of an operating altitude (i.e. nozzle exit condition). The cold gas thruster has an expansion area ratio of 60. For ejector design, the parameters like the length to diameter ratio (L/De) and ejector nozzle stagnation pressure (Ptot) seem to have significant importance, and their effects on vacuum cell pressure are studied by using both analytical and experimental techniques. Role of second throat type diffuser in the overall pressure recovery is also investigated. The required exit pressure is achieved by a two-stage ejector system having constant area mixing chambers. The ejectors’ stagnation pressures are considered from 12 to 20 bars, while (L/De) is considered from 4.5 to 15 and 2.5 to 16.5 for Ejector-1 and Ejector-2, respectively. Minimum vacuum cell pressure is achieved with the (L/De) of 7 and 9 and stagnation pressure (Ptot) of 18 and 20 bars for Ejector-1 and Ejector-2, respectively. The thrust coefficient (CF) and the specific impulse (Isp) were found in good agreement with the theoretical results.
2 citations
19 Jan 2023
TL;DR: In this paper , the authors used a two-stage ejector system to create the desired diffuser back pressure conditions at the exit of a second throat diffuser during the tests.
Abstract: In this paper, we investigate several cold gas thrusters with different nozzles geometries numerically and experimentally in the Purdue Altitude Chamber Facility where we use a diffuser-ejector system to simulate high altitude conditions. Here, the main aim is to study the effects of different thruster nozzle geometries and axial/radial gaps between the thruster nozzle exit and the diffuser inlet on the performance of the diffuser-ejector system. We used a two-stage ejector system to create the desired diffuser back pressure conditions at the exit of a second throat diffuser during the tests. Results to date focus on one of the thrusters having a highly truncated ideal contoured nozzle. While testing that cold gas thruster, we used the mass flow provided by the thruster as the suction load for the ejector system and completed its full performance characterization. We also studied the minimum starting, operating pressure ratios, and hysteresis behavior of the second throat diffuser. We obtained a noticeably small hysteresis region. We also found that the diffuser starting pressure ratio increased with increasing the axial gap between the nozzle exit and the diffuser inlet. Future work will consist of completing the tests for all the thrusters nozzle geometries. Furthermore, as we had used the same diffuser and the ejector system with a hybrid motor at simulated high altitude conditions, we compared the performance of the diffuser – ejector system with both hot and cold gas flows. By doing that, we can explore the effects of the specific heat ratio and the temperature of the exhaust gas flow on the diffuser-ejector system performance.
References
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TL;DR: In this paper, the performance of six well-known turbulence models for the study of supersonic ejectors was evaluated and the results showed that the k-omega-sst model agrees best with experiments.
Abstract: Supersonic ejectors are widely used in a range of applications such as aerospace, propulsion and refrigeration. The primary interest of this study is to set up a reliable hydrodynamics model of a supersonic ejector, which may be extended to refrigeration applications. The first part of this work evaluated the performance of six well-known turbulence models for the study of supersonic ejectors. The validation concentrated on the shock location, shock strength and the average pressure recovery prediction. Axial pressure measurements with a capillary probe performed previously [Int. J. Turbo Jet Engines 19 (2002) 71; Conference Proc., 10th Int. Symp. Flow Visuzlization, Kyoto, Japan, 2002], were compared with numerical simulations while laser tomography pictures were used to evaluate the non-mixing length. The capillary probe has been included in the numerical model and the non-mixing length has been numerically evaluated by including an additional transport equation for a passive scalar, which acted as an ideal colorant in the flow. At this point, the results show that the k–omega–sst model agrees best with experiments. In the second part, the tested model was used to reproduce the different operation modes of a supersonic ejector, ranging from on-design point to off-design. In this respect, CFD turned out to be an efficient diagnosis tool of ejector analysis (mixing, flow separation), for design, and performance optimization (optimum entrainment and recompression ratios).
323 citations
"Modeling of Two-Stage Ejector for H..." refers background in this paper
...For a second throat-type diffuser, although the second throat contraction has a strong impact, the ramp angle does not have a significant effect on the operational characteristics [14]....
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TL;DR: In this article, a straight cylindrical supersonic exhaust diffusers (SED) using cold nitrogen and hot rocket exhaust gases as driving fluids were used to evaluate the effects of the ratios of the SED area to rocket nozzle throat area (Ad/At), SED areas to rocket exhaust manifold exit area, SED length to its diameter (L/D), and specific heat ratio of the driving gases (k) on the minimum starting pressure ratio, (Po/Pa)st, of SED.
Abstract: Experiments were carried out on straight cylindrical supersonic exhaust diffusers (SED) using cold nitrogen and hot rocket exhaust gases as driving fluids, in order to evaluate the effects of the ratios of the SED area to rocket nozzle throat area (Ad/At), SED area to rocket nozzle exit area (Ad/Ae), SED length to its diameter (L/D) and specific heat ratio of the driving gases (k) on the minimum starting pressure ratio, (Po/Pa)st, of SED. The rocket nozzle and SED starting transients were also simulated in the models. The study reveals that (Po/Pa)st increases monotonically with increase in (Ad/At) and k. One-dimensional normal shock relations were used in predicting the (Po/Pa)st since the compression in long ducts is basically a normal shock process. Predicted values of (Po/Pa)st were validated with experimental data. SED efficiency factors(ηns) were arrived at based on one-dimensional normal shock relations. ηns goes down at higher values of (Ad/Ae). (Po/Pa)st is lower for lower k values for the same (Ad/At). Cylindrical SEDs exhibit no hysteresis. The results of this investigation were utilised in validating the design of high altitude test (HAT) facility for testing the third stage motor (PS-3) of Polar Satellite Launch Vehicle (PSLV). The simulation of starting transients in the model revealed that the HAT facility shall not be operated in the unstarted phase, because the rocket nozzle may fail due to violent oscillations of the vacuum chamber pressure. These experimental data were also utilised for designing a SED for PS-3 sub-scale motor, the results of which are covered in this paper. The accuracy of measurements are within a range of ±0.4%. Error analysis of the data were carried out and are presented in Appendix A .
64 citations
"Modeling of Two-Stage Ejector for H..." refers background in this paper
...The main concern in designing a SED [3–5] is to keep the HAT facility length and starting pressure ratio as low as possible....
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TL;DR: In this paper, a second-order high-resolution scheme for solving the new Lagrangian Euler equations is employed to accurately resolve the complicated shock patterns and associated slip lines and their interactions.
Abstract: A computational analysis of the two-dimensional supersonic inviscid flowfield in a second-throat ejector-diffuser (STED) system is presented. A second-order high-resolution scheme for solving the new Lagrangian Euler equations is employed to accurately resolve the complicated shock patterns and associated slip lines and their interactions. A parametric study covering a variety of Xst and Ost is implemented to investigate their effects on the flow structure in STED as well as its performance. Results suggest that the averaged Mach number along the entrance plane of the second throat is a suitable criterion for the justification of the performance of STED. With this criterion, an optimal design insuring the largest pressure recovery can be achieved.
41 citations
"Modeling of Two-Stage Ejector for H..." refers background in this paper
...An ejector–diffuser system is said to be “started” when the area-averagedMach number along the entrance plane of the mixer throat is supersonic; in fact, this criterion has been found suitable for accepting the performance of STED [6]....
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TL;DR: In this article, the starting transient and plume blowback at diffuser breakdown of a straight cylindrical supersonic exhaust diffuser with no externally supplied secondary flow are numerically investigated.
Abstract: Starting transient and plume blowback at diffuser breakdown of a straight cylindrical supersonic exhaust diffuser with no externally supplied secondary flow are numerically investigated. Fluctuating pressures occurring in the transitional periods are measured by a small-scale cold-gas simulator for high altitudes. The flow-fields evolving in the diffuser-type ejector are solved by a preconditioned Favre-averaged Navier–Stokes equations with a low-Reynolds number k–e turbulence model edited for comprehending compressibility effects. In the steady operation regime, strong momentum transfer by a high-speed jet at diffuser inlet is balanced while it forms a recirculation zone at the diffuser inlet. Flow separation occurs at the diffuser downstream during the starting transient of the diffuser. During the starting transient and diffuser breakdown, abrupt changes in internal shock structures including a transition from Mach to regular shock wave reflection occur. Diffuser choking is advanced by anchoring the sonic line from the nozzle onto the wall of diffuser inlet before the reversed flow from the downstream subsonic region reaches the upstream of the diffuser. Immediately after the diffuser breakdown, higher ambient pressure acts on the upstream flows and the exhaust gases from the nozzle surge into the vacuum chamber. Confronted with the plume blowback, the exhaust flows are blocked and non-monotonic pressure recovery occurs.
36 citations
"Modeling of Two-Stage Ejector for H..." refers background in this paper
...The main concern in designing a SED [3–5] is to keep the HAT facility length and starting pressure ratio as low as possible....
[...]
TL;DR: In this article, the performance of a supersonic air ejector in the case of zero-secondary flow was studied by varying the Mach number of the primary nozzle and the throat area ratio of the mixing tube to the primary.
Abstract: The performance of a supersonic air ejector in the case of zero-secondary flow was studied by varying the Mach number of the primary nozzle and the throat area ratio of the mixing tube to the primary nozzle. As the result, it has been shown that an optimum throat area ratio exists for each Mach number of the primary nozzle in which case the vacuum performance of the ejector becomes maximum with a minimum stagnation pressure of the primary flow. The physical meanings of the optimum throat area ratio and other ejector characteristics have been clarified from the results of pressure measurements and optical observations. Furthermore, modifying a previous method in the case of a constant-area mixing tube, the ejector performance was calculated for a variable-area mixing tube in the case of a larger throat area ratio than the optimum one. Calculated values agree well with the experimental results.
35 citations