Nematollah Fouladi

The ratio of diffuser inlet to nozzle exit area (Ad/Ae) is one of the most important parameters in designing and evaluating the performance of a high-altitude test facility (HATF). Typically, at motor pressures near start or operating pressure of the diffuser, the vacuum chamber pressure oscillates at a high Ad/Ae, which may be dangerous for the HATF. In the present study, the effect of the Ad/Ae on the starting and breakdown performance of a second throat exhaust diffuser (STED) has been investigated experimentally using a HATF supplied by high-pressure cold air apparatus. Upon examining the overall performance of the diffuser, it is observed that with an increase in Ad/Ae, the oblique sealing shock wave at the diffuser inlet undergoes significant changes in location and strength. This leads to a further reduction of the flow total pressure and a stronger separation of the flow along the diffuser. Consequently, flow instability arises during starting or breakdown phases near the STED starting or operating pressure. By frequency analysis, 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 in these transient situations becomes larger. Additionally, there is a significant decrease in nozzle pressure related to the onset of pressure oscillations. Furthermore, with an increase in Ad/Ae from 1.27 to 7.81, the minimum starting pressure and breakdown pressure of STED increase by 20.33% and 28.47%, respectively. Also, the hysteresis range in performance of STED vanishes at high Ad/Ae values.

This research investigates the process of starting and breakdown of second throat diffuser during high-altitude test of conical nozzle with a high expansion ratio. The subscale experimental setup includes a conical nozzle with an expansion ratio of 53, plus a second throat diffuser with a contraction ratio of 1.85, using compressed air as the working fluid. Numerical simulation has been employed to identify the flow physics during the unsteady process of diffuser startup and breakdown. According to the results, the starting process is divided into three distinct stages. In the first stage, the exhausted flow from the nozzle enters the vacuum chamber, leading to an increase in vacuum chamber pressure. The second stage, corresponding to the period before full flow establishment in the conical nozzle, exhibits a relatively constant slope. In the final stage, associated with the transition from over-expanded to under-expanded states, the slope of the rate of evacuation development descends. Next, the vacuum degradation in termination process has been analyzed, and it has been found to include three stages: high slope, middle slope, and low slope. The flow physics during the start process, similar to the results observed in other conical nozzles, exhibits only a Mach reflection structure. However, during the termination process, the flow physics involve a combination of both structures, including Mach reflection and Cap Shock. The results indicate that during the start, an internal shock only interacts with the separation shock, and no special change occurs in the Mach reflection structure. In contrast, during the termination process, unlike what has been reported in previous studies on conical nozzles, the structure of cap shock waves and restricted shock separation patterns are also observed. Another distinction between the starting and termination processes is related to the pressure distribution in the diffuser wall. The wall pressure at the diffuser inlet during the termination process has been reported to be 90% higher than during the starting process.

One of the frequently reported defects of RANS-based turbulence models is overestimation of turbulent kinetic energy production in high speed separated flow problems, which causes significant prediction errors. The correct estimation of such flow in thrust optimized parabolic nozzles extremely depends upon the accurate prediction of the onset of flow separation. In this paper, firstly, the significant error of conventional RANS-based turbulence models is shown to predict the onset of flow separation in this type of nozzles. Then, the prediction accuracy is improved through the modification of the essential parameters of the generalized k-ω (GEKO) turbulence model. It was found that modifying the separation and mixing parameters of the GEKO model to realize the turbulent kinetic energy production resulted in the accurate prediction of onset of flow separation at the extensive range of nozzle pressure ratios. Using this modified model with new coefficients reduced the error of about 30% of the k-ω-sst model in estimating the onset of flow separation. Also, the nozzle pressure value at which the transition from free shock separation (FSS) to restricted shock separation (RSS) occurs is well predicted by this approach. After strengthening the turbulence model, the flow physics has been investigated with increasing and decreasing nozzle chamber pressure. The length of the separation shock and reflected shock waves which impose the presence of FSS or RSS patterns and transitional phenomena are discussed. Our new findings show that unlike the transition from FSS to RSS, the inverse transition from RSS to FSS did not depend on the length of the separation and reflective shocks.

This study presents a combined numerical and experimental investigation of the starting process of a second throat diffuser during ground testing of a thrust-optimized parabolic (TOP). In this investigation, compressed air has been utilized as the operating fluid in a subscale experimental setup. The study examines three scenarios with varying nozzle pressure profile, including two cases of start and one case of un-start. Additionally, this study employs numerical simulations to identify and analyze the physical phenomena that occur at each stage of the start and un-start processes in these cases. The results for the case started at a relatively low nozzle pressure profile (24.5 bar max) indicate that the vacuum generation process during high-altitude testing of TOP nozzles can be broken down into five stages. The first stage involves an increase in pressure within the vacuum chamber during the early moments of the starting process. In the second stage, vacuum generation occurs gradually as the nozzle operates under free shock separation (FSS). This is followed with the reappearance of small fluctuations in the vacuum chamber pressure due to transition from the FSS to restricted shock separation (RSS) flow pattern (third stage). The fourth stage begins with the predominance of the shock separation and recirculation (SSR) flow pattern inside the nozzle, resulting in gradual vacuum generation. This stage terminates upon transformation of the cap shock structure into a regular reflection structure. In the final stage of vacuum generation, the evacuation rate is almost half of the fourth stage. This is attributed to the establishment of expanded and under-expanded conditions, as well as the impingement of the nozzle outflow jet with the wall and the onset of start conditions. Next, the results of vacuum generation have been studied at higher nozzle pressures profile (34 bar max). The results indicate that increasing the nozzle pressure rate not only reduced the starting time by 23%, but also significantly reduced the pressure fluctuations in the evacuation process. In fact, at higher nozzle pressure, the third stage is almost eliminated. In the un-started case, where the nozzle pressure is lower than the minimum starting pressure, fluctuations occur in the vacuum chamber pressure due to the dynamics of the diffuser inlet recirculation bubble and the transition of the nozzle separation pattern from RSS to SSR and vice versa.

High-altitude test facilities are usually used to evaluate the performance of space mission engines. The supersonic exhaust diffuser, a main part of high-altitude test facility, provides the required test cell vacuum conditions by self-pumping the nozzle exhaust gases to the atmosphere. However, the plume temperature is often much higher than the temperature the diffuser structure is able to withstand, usually above 2500 K. In this study, an efficient cooling system is designed and analyzed to resolve the thermal problem. A water spray cooling technique is preferred among various existing techniques. Here, a new algorithm is developed for a spray cooling system for a supersonic exhaust diffuser. This algorithm uses a series of experimental and geometrical relationships to resize the governing parameters and remove the required heat flux from the diffuser surface. The efficiency of the newly designed cooling system is evaluated via numerical simulations. The utilized numerical technique is based on the discrete-phase method. Various computational studies are accomplished to enhance the accuracy of numerical prediction and validation. The present numerical study is validated using experimental results. The results show that the realizable k-ɛ method is superior compared to other Reynolds-averaged Navier–Stokes models.