nematollah fouladi

One of the most important parameters in the operation of the diffuser of the vacuum simulator is the ratio of the cross-sectional area of the diffuser inlet to the outlet cross-section of the nozzle. Increasing this parameter increases the expansion length of the supersonic flow and causes instability in the process of starting up the diffuser. In the present study, the effect of the different diffuser inlet to nozzle exit area ratio (Ad/Ae) on the starting performance of a second throat exhaust diffuser has been investigated experimentally. Vacuum simulation tests have been carried out on several conical nozzles with different exit area by a high-pressure air apparatus. Investigations show that at high Ad/Ae before starting the diffuser, oscillations in the pressure of the test chamber are created due to the instability of the flow at the diffuser inlet. By frequency analysis of the vacuum chamber pressure, it is observed that as Ad/Ae increases, the number of oscillatory modes of the vacuum chamber pressure increases and the dominant frequency of the oscillations before starting the diffuser becomes larger. It was also observed that, generally, with the increase of Ad/Ae the amount of motor pressure corresponding to the start of oscillations decreases. In addition, according to the investigations, with the increase of Ad/Ae from 1.27 to 7.81, the minimum starting pressure of second throat exhaust diffuser increases by 20.33 %.

The purpose of this research is to evaluate a ground test bed of an orbital transmission engine with pre-evacuation of the engine's internal space. In the usual tests on the ground, the initial pressure of the engine is atmospheric pressure. While during the orbital mission, the internal space of the engine may be in the vacuum pressure. Therefore, it is necessary to test the performance by pre-evacuating its internal space. In this research, the suitability of an exhaust diffuser for this type of test is investigated numerically. The unsteady numerical simulations have been done by applying the pressure-time profiles of the engine as the boundary condition of the inlet pressure. Investigations show that the two phenomena of flow being supersonic in the diffuser at very low engine pressures and the discharge of the return flow to the vacuum chamber prevent the significant influence of environmental conditions on the flow inside the nozzle. So, from the initial moment to the stable working of the diffuser, the flow in the first half of the nozzle is in the supersonic state. Therefore, the internal ballistics of the engine is evaluated independently of the conditions of the outside environment.

In the present research, the experimental and numerical investigation of the starting process of the second throat diffuser with a parabolic nozzle containing expansion ratio of 35 has been conducted at a relatively low total pressure. This investigation uses an experimental setup known as a small-scale high-altitude test facility with compressed air as the working fluid. Using numerical simulation, the physical phenomena occurring in each stage of vacuum generation have been identified and analyzed. The results show that the process of vacuum generation in the high-altitude simulator of the second throat type with the presence of Thrust Optimum Parabolic (TOP) nozzles included four stages. In the first stage, the vacuum was gradually created by using the nozzle in the free shock separation (FSS). In the second stage, which started as soon as the transition from the FSS pattern to restricted shock separation (RSS), the vacuum generation was slow and accompanied by oscillation. Vacuum generation was gradual in the third stage, which corresponds to the beginning of the shock separation with recirculation(SSR) and continues until the end of regular reflection (RR). And finally, the last stage also coincides with the structure of expanded, under-expanded conditions, the impact of the jet exiting the nozzle with the diffuser wall, and the establishment of start-up conditions, creating a vacuum at a slower rate than in other stages.

The aim of this study is to evaluate the performance of water-jacket cooling system for thermal protection of an exhaust large dimension diffuser at high heat fluxes in a wide range of coolant pressure. For this purpose, a suitable design algorithm has been developed. Using the present method, the parameters of the water-jacket cooling system are determined so that in addition to satisfying the temperature conditions of the metal body, the total pressure drop has remained in the desired range. In the following, the capability of numerical code to design and performance analysis of the cooling system has been evaluated in coolant pressure of 3 to 50 bar and high heat fluxes up to 3.5 MW/m2. The present studies show that the proper selection of coolant pressure is very important in the design of the cooling system with optimal mass flow rate and minimum coolant dimensions, especially at high heat fluxes, so that increasing the coolant pressure from 3 to 10 bar, in addition to significantly reducing the dimensions of the cooling system, reduces the mass flow rate by 75%.

Complex flow separation in thrust optimized parabolic nozzles in the over-expanded condition is one of the challenging issues of many numerical investigations. The correct estimation of a thrust optimized parabolic nozzle performance extremely depends upon the accurate estimation of the onset of flow separation. Literature review indicates that conventional Reynolds-averaged Navier–Stokes turbulence models have a significant error in predicting the onset of flow separation in these types of nozzles due to the overestimating of turbulent kinetic energy production. Recently proposed generalized k-omega has made it possible to rectify numerical simulations based on governing physics and using limited experimental results. In the present study, the flow physics in the LEA_TOC nozzle has been investigated with the numerical simulation approach. At the first, the significant error of conventional Reynolds-averaged Navier–Stokes turbulence models is shown to simulate flow separation in this type of problem. Then, the generalized k-omega parameters are modified based on the limited experimental result of the LEA_TOC nozzle, and the ability of this model has been evaluated to estimate the flow physics under different pressure ratios. Numerical investigations show that generalized k-omega has a high capability for accurately estimating the onset of flow separation at a wide range of nozzle pressure ratios. Applying the corrected generalized k-omega has resulted in an improvement of about 30% in the estimation of the onset of separation in the over-expanded LEA_TOC nozzle compared to the k-ω-SST model.

In the present study, the effect of pre-evacuation on starting process of a second throat exhaust diffuser has been investigated experimentally by examining a thrust optimized parabolic nozzle. An experimental setup called high-altitude test facility has been used with compressed air as operating fluid. According to the importance of area ratio parameter (Ad/Ast) of a second throat diffuser, the effect of this parameter variation has been examined on the start-up performance of the nozzle and diffuser. In all of the diffuser geometries, to evaluate the instantaneous performances, the pressure in the nozzle chamber has charged instantly in two modes; with and without pre-evacuation. then, the vacuum chamber pressure and static pressure distribution along the diffuser are measured by a data acquisition system. The results show that pre-evacuation in the test chamber reduces the start-up time of the diffuser by 50 to 60%. In addition, pre-evacuating the test chamber eliminates the destructive transition phenomenon from the flow separation pattern during start-up of the nozzle and diffuser. Also, It is observed that with the narrowing of the diffuser’s second throat duct, the minimum starting pressure of the diffuser increases and eventually flow chocks at the second throat in a certain area ratio.

This paper is presented to investigate the deposition effect on a second throat exhaust diffuser performance. In the numerical simulation, the blockage of the diffuser due to the deposition of aluminum oxide is considered by a gradual and time-dependent cross-section constriction. In the initial conditions, the supersonic flow has been established in the nozzle and diffuser. Diffuser cross-section area is reduced by using a dynamic mesh method during the solution. The flow is considered compressible, viscous, and 2 dimensional axis-symmetric. The k-ω shear stress transport turbulence model is used to solve the turbulent flow field. Diffuser blockage (n) is equal to the ratio of instantaneous and primary diameters of the second-throat. By changing the value of n from 1 to 0.75, the onset of flow separation is moved to the downstream location in the diffuser. This results in a considerable reduction of total pressure loss and then improves the flow pressure recovery. Decreasing parameter n from 0.75 to 0.64, the flow structure is subjected to severe changes and the separation of the flow occurs near the diffuser inlet or inside the nozzle. In this condition, the diffuser state changes from starting to un-starting mode. Therefore, the vacuum condition vanishes in the test chamber.

Usually, ground testing of space engines is performed in high altitude test facility. The facility is equipped with a supersonic diffuser that expels automatically engine gases to the atmospheric pressure and maintains a vacuum pressure around its nozzle and motor. If the motor pressure is lower than a certain amount, the supersonic flow in the diffuser will not be established. In this situation, it is necessary to use auxiliary ejector at the end of the diffuser. In the present study, a new algorithm has been developed in the design of supersonic ejector. Unlike conventional methods, this algorithm can be used for different primary (input from the ejector nozzle) and secondary fluids (outlet from the diffuser end). In the design of the ejector, the main parameters are determined by the algorithm, while the secondary parameters are selected from the empirical test results of the proved references. In this algorithm, a safe margin is considered for the safe operation of the ejector and this margin is predicted by numerical simulation. Also, numerical simulation is used to validate the design method. Finally, using the proposed algorithm, an ejector is designed to reduce the start-up pressure of an engine-diffuser assembly. An integrated simulation of the diffuser-ejector was performed for both cases that rocket motor is off and on, and the appropriateness of the designed ejector was confirmed in both modes of operation

A supersonic exhaust diffuser provides the required test cell vacuum conditions by self-pumping of nozzle exhaust gases to the atmosphere in high altitude simulator. However, the plume temperature is often much higher than the allowable temperature of the diffuser structure. Moreover, in solid-fuel engines, the impact of aluminum oxide particles on the diffuser wall internal surfaces with high temperatures intensifies critical thermal conditions. In the present study, a spray cooling system design method is presented for a supersonic exhaust diffuser. The method is evaluated by performing several experimental tests. First, in order to identify the critical temperature region, the test of the motor with a chamber pressure of 60 bar and a chamber temperature of 3100 ° C is performed with a non-cooled metal diffuser. The results indicate that the temperature of the diffuser body in the inlet and ramp regions reaches a temperature above 1500 ° C, which leads to the melting and perforation of the diffuser in these regions. Two other tests are performed with average motor chamber pressures of 33 bar and 55 bar along with the spray cooling of diffuser body. The results show that the designed cooling system keeps the maximum temperatures of the external surface of the diffuser at the values smaller than 200 and 400 ° C in these tests. The achieved critical temperatures are well matched with the respected ones in the design procedure. This confirms the present spray cooling system design procedure.

In this research, the performance study of an altitude test simulator at transient phase of motor burnout is conducted by numerical approach. Using a time dependent pressure profile of a motor in burnout phase, the unsteady exhaust flow is simulated in the main sections of the altitude test simulator, i.e. high expansion ratio nozzle, second throat exhaust diffuser, and vacuum chamber. Present investigation shows that in spite of the high pressure starting condition of the altitude simulator, the supersonic flow in the diffuser tends to breakdown at relatively low combustion pressure in the motor terminating phase. At the breakdown condition, the nozzle exhaust hot gases directed into the vacuum chamber through the annular gap between nozzle and diffuser walls. Present simulation shows that the overall temperature of fluid in vacuum chamber is reached up to 2000 K after the motor burnout. Really, It is potentially dangerous for measurement instruments inside the vacuum chamber. Furthermore; in this research, the influences of backflow arrester (BFA) in variations of vacuum chamber temperature are studied at this transient terminating phase. It has been shown that, the BFA can decrease the exhaust streaming into the chamber in initial breakdown periods. Results show that utilizing a suitable BFA size, the safe temperature condition could be established in vacuum chamber during and after the transient terminating phase of motor.

In the present research about 50 high altitude test facilities (HATFs) in the world and ejectors of them have been analyzed. Practical information such as type of the ejector, operating fluid, procedure of fluid generation, operating pressure and temperature, Thermal protection and costs of construction and operation can be useful in developing space industries. The present study shows that in 87% of the 50 HATFs, central ejectors and in 76% of them water steam have been used. Water spray cooling is often used to provide thermal protection for ejectors. Based on the available references, the water mass flow rate for cooling system is in the range of 1.8 to 3.9 times of the mass flow rate of exhaust gases of the engines with thrust in the range of 0.11 to 204 kN. To produce water steam, the HATFs for solid and liquid fuel engines frequently employ the steam accumulators and steam generators, respectively.

During transient phase of motor burning in an altitude test simulator, the low momentum exhaust combustion gases could not establish supersonic flow in the exhaust diffuser. This leads to a reverse flow of the exhaust gases into the vacuum chamber through the annular gap between nozzle and diffuser walls. This spoils the vacuum in the chamber and it is potentially dangerous for measurement instruments inside the vacuum chamber. In this research; at the first, the physics of the flow at initial transient phase of motor burning is investigated numerically and the backflow phenomenon is illustrated. Then, the influences of backflow arrester (BFA) in variations of vacuum chamber pressure and temperature are studied at the transient starting phase. It has been found that applying BFA to decrease the backflow to the vacuum chamber has no significant effect on starting time of the diffuser but it has strong effect on peak pressure and temperature reduction of the vacuum chamber. It is although found that attaining to the steady condition in vacuum chamber is delayed with this instrument. However, using one-way backflow arrester of reasonable size, both the starting time of diffuser and the peak temperature of the vacuum chamber experience a strong reduction. Although, the vacuum chamber steady condition is accelerated with this alternative. Therefore, this device is more suitable than conventional BFA for transient starting phase of the altitude test simulators.

In this paper, the design algorithm of a second throat exhaust diffuser applicable in altitude tests of large expansion ratio nozzles is presented. In this algorithm, the geometric parameters of the exhaust diffuser are classified into primary and secondary parts. The primary geometric parameters are calculated from normal shock theory incorporating with a correction coefficient. However, the secondary parameters are selected from the previously reported experimental results. Numerical simulation tool is utilized to satisfy the design candidates and to finalize the correction factor. Axis-symmetric compressible Navier–Stokes equations incorporated with two equation Kω-SST turbulence model are solved to extract the supersonic exhaust diffuser flow features. As a first stage of numerical analysis, we use an unsteady pressure-based solver to accelerate the solution procedure. At the second stage, we use steady density-based solver to enhance the accuracy of our solutions. The current numerical method is properly validated by experimental reported results in the literature. Finally, we focused on simulation results of a designed diffuser and described the flow futures at different boundary conditions. The simulation results are confirmed that the designed diffuser is suitable for proposed altitude test.