High-altitude Test Research Articles

Experimental evaluation of the influence of the diffuser inlet to nozzle exit cross sectional area ratio on pressure oscillation in a high-altitude test facility

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.

Fault tree analysis for risk assessment of rocket engine testing process at high altitude test facility

Fault tree analysis for risk assessment of rocket engine testing process at high altitude test facility

Gas dynamics at starting and terminating phase of a supersonic exhaust diffuser with a conical nozzle

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.

Numerical Investigations At High Altitudes For Nozzles Performance Monitoring

Large area ratio nozzles that are employed for space propulsion applications are not meant to be tested at the ground-level conditions because of the flow conditions that occurs in their divergent section due to their low exhaust pressures. Therefore, it is desired to perform the experimental evaluation at high-altitude to evaluate the performance of the said nozzles. Generally, a high-altitude test facility, consisting of a supersonic exhaust diffuser, is generally employed. In thispaper, a second-throat exhaust diffuser has been numerically investigated to predict its minimum starting pressure, and to understand the flow physics in the diffuser using Computational Fluid Dynamics (CFD). Numerical computation of the flow field in the diffuser system is done for the two cases: 1) when the nozzle and the diffuser system are initially evacuated to a low pressure, and 2) when the nozzle and the diffuser system are initially at ambient pressure (1 bar). Simulations have been carried out for cold flow situation (Ȗ=1.4) over a range of nozzle inlet stagnation pressures. Numerical results compare favorably with the theoretical and experimental results and provide adequate insight to the flow physics and internal shock structures formed in thediffuser in addition to minimum starting pressure, diffuser wall pressure and vacuum thrust.

Starting transient analysis of second throat exhaust diffuser in high-altitude test of a thrust optimized parabolic nozzle

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.

Experimental Study On Bird Streamer Flashover for V-Type Insulators of DC Transmission Line in High Altitude Area

In order to obtain the bird streamer flashover characteristics in high altitude areas, the discharge of DC V-type string insulators was studied in high-altitude test base at 4300 m in Tibet. The study results show that medium viscosity bird streamer is easy to form a continuous channel and cause flashover. The flashover probability of positive polarity transmission lines is slightly higher than that of negative. The conductivity of bird excrement almost has no effect on the flashover probability in the range of 1.9∼20.9 mS/cm. To increase the vertical air gap distance can reduce bird streamer flashover probability. Based on the experiment results, the recommended vertical air gap distance values of ±400 kV DC transmission line were proposed. Considering the influence of wind and bird streamer falling angle, the protection radius was set 3.5 m at 3500∼5000 m altitude area. Bird streamer flashover is a mixed discharge of air gap breakdown and the surface discharge along the bird streamer channel, at the altitude of 4300 m area, the minimum breakdown potential gradient is about 2.1 kV/cm, and the discharge between bird streamer and shielding ring can be approximately equivalent to rod-rod gap discharge.

Optimal diffuser geometry design and verification for micro-scale high altitude test stand

Optimal diffuser geometry design and verification for micro-scale high altitude test stand

Multifidelity Simulation Research on the Low Reynolds Number Effect on the Engine Performance at Different Altitudes

Abstract With the rapid development of unmanned aerial vehicles, the effect of the low Reynolds number on gas turbine performance and high-altitude endurance has received extensive attention. However, the existing three-dimensional component modeling cannot meet the design requirements of the whole engine level, and the accuracy and physical principles of the existing engine empirical correction cannot be guaranteed. Through the study of a single-shaft turbojet engine, this paper adopts a versatile and accurate coupling method, which combines the volume method and the full coupling method and conducts multifidelity simulation research on the zero-dimensional engine model and the three-dimensional component model. Then, based on the high-altitude test data, under typical operating conditions, compared with the existing empirical correction method in gasturb, the accuracy of the engine inlet flow, fuel flow, thrust, and exhaust gas temperature predicted by the volume-based fully coupled method is improved by 6.2%, 7.9%, 4.7%, and 11.4%, respectively. Next, as the flight altitude rises from 0 km to 21 km, the Reynolds number reduces, the working lines approach the surge lines, and the maximum mass flow rate and the efficiency of the engine components gradually decrease. In addition, in the flow field of the components, the thickness of the boundary layer increases, the shock wave intensity decreases, and the position moves forward. The core innovation of this article is that it provides a creative multifidelity evaluation method for gas turbines to effectively solve the problems of insufficient accuracy of the existing empirical correction methods and the inability of the component design to meet the gas turbine requirements in the study of the low Reynolds number effect at different altitudes, which significantly strengthens the connection among the component internal flow field information extraction, the component characteristics analysis, and the gas turbine performance matching. Moreover, it is conducive to the scientific design of the advanced unmanned aerial vehicles' power.

Investigation of nozzle flow in high altitude test facility

A high altitude test facility was developed for the experimental studies on nozzles for various levels of vacuum. The current study is focused on the performance of the nozzle under various altitude condition and to characterized the high altitude test facility. A supersonic nozzle designed for Mach 2.5 is used for the study. Compressed air is taped from the high pressure plenum having a pressure of 20 bar and is regulated and expanded through the nozzle. The inlet pressures for the study is varied from 4.5 to 10 bar. The nozzle is within the enclosure which is evacuated to 0.7–0.02 bar. Schlieren is used to view the flow condition at the end of the nozzle. A nozzle for 2.5 Mach is designed and tested in HAT facility. The nozzle design is validated with the CFD for various NPR. The high altitude test facility is characterized for various NPR and is found to be optimum flow at 14 NPR for 33 s at an inlet pressure of 4.5.

Research on the volume-based fully coupled method of the multi-fidelity engine simulation

Numerical simulation is an indispensable means to develop advanced aero engines. To improve the simulation accuracy of aero engines under limited computing resources, coupling methods are commonly used to investigate the multi-fidelity engine simulation. However, for complex high-precision component models, the obvious problem in the application of existing methods is reflected in the high computational cost. Through the investigation of a single-spool split-flow turbofan engine, this paper presents a rapid and accurate coupling method to research the multi-fidelity simulation of three-dimensional component models and zero-dimensional engine model, which creatively combines the advantages of the volume method and the fully coupled method. Then, through numerical examples of the turbofan engine and another single-spool turbojet engine, the volume-based fully coupled method is compared with the existing coupling methods in terms of accuracy, computational cost, and simplicity. Moreover, the experimental data of the ground test and high-altitude test are performed to verify the effectiveness of the coupling methods. And the results indicate that the volume-based fully coupled method matches well with the test data. For the numerical examples of the turbofan and turbojet engines, the volume-based fully coupled method consumes 56.74 and 39.17 hours to realize the multi-fidelity simulation respectively. And the former example reduces the computational burden by 52.76% and 38.12% respectively compared with the existing iterative coupled method and fully coupled method. While the latter example saves the computational cost by 50.11% and 36.28%. The main novelty of this paper lies in that it affords an effective solution capable of significantly decreasing the computational burden of the multi-fidelity simulation model based on guaranteeing simulation accuracy. Moreover, it fully embodies the design idea of the whole engine guiding the engine components, and it could provide more realistic requirements for component design during the design phase of future new concept aero engines in the absence of component characteristics.

Starting transients in second throat vacuum ejectors for high altitude testing facilities

The present study aims to investigate the process of vacuum generation in a second throat vacuum ejector system employed for the high altitude testing of rocket motors. The study shows that the vacuum generation progresses with four distinct stages during the initial start-up process. The initial stage consists of a gradual evacuation process (first stage) which is followed by a transition region (second stage) in which the pressure starts to drop fastly and eventually leading to a rapid evacuation stage (third stage). This is followed with the reappearance of a gradual evacuation stage (fourth stage) which terminates with the onset of the started mode operation. It is found that the various stages of evacuation are closely linked to the internal shock wave movement in the supersonic nozzle. The gradual evacuation is found to be caused by the relatively large nozzle exit pressure caused by the multiple shock cells in the nozzle. The rapid evacuation occurs due to the sudden movement of the first shock cell to the nozzle exit and the reappearance of the gradual evacuation happens during the transformation of the nozzle jet from an overexpanded to underexpanded state. The parametric studies show that with increase in diffuser convergent cone angle, the contribution of vacuum generation from the fourth stage reduces due to a reduction in the maximum possible jet expansion. A same trend in vacuum generation is observed when nozzle distance from the second throat section of the ejector diffuser reduces.

Computational and experimental study of the influence of the shape of nozzle supersonic part on the flow structure in the gas-dynamic flow path of a model high-altitude test facility

Computational and experimental study of the influence of the shape of nozzle supersonic part on the flow structure in the gas-dynamic flow path of a model high-altitude test facility

Computations of Side Loads in a Thrust-Optimized Parabolic Nozzle During High-Altitude Testing

The present work is to computationally analyze side loads in a thrust-optimized parabolic nozzle during high-altitude testing. The engine startup and shutdown processes of the Korean Space Launch Vehicle II third-stage rocket engine combustor are examined using three-dimensional time-accurate computations. The generation of side loads is closely related to the transition of the separation pattern due to the nozzle shape. However, the results also show that the side loads are significantly affected by the unique flow features of high-altitude testing, such as the flow entering through the gap between the nozzle and the diffuser, as well as the pressure waves produced by the impact of the diffuser flow breakdown. During the engine startup process, the interaction between the nozzle flow and the entrained flow produces unexpected side loads before the transition of the separation pattern. The first transition process generates two side-load peaks. The second transition does not produce any side loads owing to the smooth transition process, which is caused by the entrained flow acting like a nozzle extension. During the engine shutdown process, the initial pressure wave from the diffuser causes fluctuating side loads. The second pressure wave leads to the transition of the separation pattern, which temporarily increases the magnitude of the side loads.

Design and realization of thrust measurement system for high altitude testing of cryogenic rocket engine

Design and realization of thrust measurement system for high altitude testing of cryogenic rocket engine

Computer Simulation Research on Flow Field of SRM High Altitude Simulation Test

The simulation test of passive ejection type high-altitude solid rocket motor (SRM) was simulated by computer. For SRM passive ejector type high-altitude test, the tempering in the process of engine ignition and flameout poses a serious threat to the nozzle of the engine test device and circuit arrangement in the upper compartment. Using the N-S equation and Spalart-Allmaras model: numerical simulation on the flow structure of engine pressurize, decompression and stratochamber air phase in the testing process, and flow field of the diffuser and the cabin altitude simulation, analyzing the flow field structure of high-altitude experiment stages. The results showed the rough coincidence between the simulation curve and the measured data, with slight difference at the initial stage of air supply, the measured data and the simulation data reach the overall high degree of coincidence. During the process of building a compression and decompression, the adiabatic measures should be taken to prevent high temperature gas from damaging the engine, especially the nozzle and the head ring, in addition, the reasonable design of the nozzle contour surface can control gas separation tempering, reducing the influence caused by tempering. Stability of the flow field in the high altitude chamber of the pressure will not affect the engine.

Onset of detonation through a cold flow analysis using metallic diaphragms – an experimental approach

Modern era is looking for high speed transportation systems across globe with due comfort. Innovative design in both aerodynamic and propulsion studies have been made for achieving supersonic to hypersonic speeds ranging a claim of 1km/s under stratosphere limit. The complexity in designing such vehicles in a high-altitude test facility is borne due to the fact that simulation of such conditions as low density, no shock conditions, high rate of dissipation and shock flow comes from unavailability of sufficient technology to pressurize gases to the optimal testing conditions. The prominent ways of constructing a shock tube either involve a blow down type or a vacuum suction type. Here a blow down type shock tube require a built-up pressure comparable to the velocity at the test section. Fibre diaphragms blow out at less critical pressure and are unsuitable for high speed flows. Metal foil are incepted as a concept of diaphragms for this particular problem. Aluminium foils are stacked across a mounting flange on a settling chamber to determine the maximum pressure at which blow down occurs. The results are tabulated for an incremental pressure schema of 1 bar to 10 bar with which it was found that the wave pressure development occurs during each batch of aluminium and recorded to be between 3.5 bar and 9 bar the successive velocities at the exit was found to be increasing with increase of blast pressure.

Study on prediction model of diesel engine with regulated two-stage turbocharging system based on hybrid genetic algorithm-particle swarm optimization method at different altitudes

ABSTRACT Aiming at the problem of complicated mechanism modeling for regulated two-stage turbocharging (RTST) system of diesel engine, a hybrid genetic algorithm–particle swarm optimization (GA-PSO) method was applied to optimize back propagation neural network (BPNN) performance prediction model of diesel engine. First, using a high-altitude (low-pressure) simulation test bench for diesel engines, a full-load performance test of RTST diesel engine at different altitudes was carried out, and the regulation rules of diesel engine intake characteristics and two-stage turbocharger (TST) matching characteristics with altitude were obtained. Based on the test data, the BPNN is used to establish the boost pressure prediction model of the RTST system of the diesel engine. For the problem of slow convergence and low prediction accuracy of the BPNN, adaptive improvement measures such as momentum and learning rate are introduced, and the GA-PSO optimizes the weights and thresholds of BPNN. The performance comparison shows that hybrid GA-PSO is superior to particle swarm optimization (PSO) and genetic algorithm (GA) in terms of evolution speed and convergence accuracy. Genetic algorithm–particle swarm optimization–back propagation neural network (GA-PSO-BPNN) model has higher prediction accuracy and higher accuracy than particle swarm optimization–back propagation neural network (PSO-BPNN) model and genetic algorithm–back propagation neural network (GA-BPNN) model. Compared with the original model, the root-mean-square error of the model is reduced by 50%, the mean error percentage (MEP) is reduced by 38%, and the sum of the squared errors is reduced by 79%. The performance verification of the GA-PSO-BPNN model shows that the MEP of the model prediction results is 5.5%, and the GA-PSO-BPNN model has higher prediction accuracy.

Performance Evaluation of Second-Throat Diffuser for High Altitude Test facility

Aimed at the improvement of ability of the high altitude test (HAT) facility for rocket engine, a new designed secondary-throat diffuser with ejection of low total pressure and supersonic cold air is proposed. The main structural characteristic of the new-designed diffuser is that: a groove is added to the start of the secondary throat of diffuser, and the supersonic cold air with low total pressure is ejected into diffuser to form the gas film, which leads to a reduction of the wall temperature and thermal load of the new-designed diffuser, and an improvement of the ability of diffuser holding shock train. The effects of cold air on flow characteristics and performance of diffuser are numerically investigated in detail under different conditions including total pressure and Mach number. The results show that the wall temperature along the diffuser increases with an increase of Mach number, and decreases with an increase of total pressure; with an increase of Mach number, the starting position of pressure rise moves upstream to the inlet of diffuser; and with an increase of total pressure, the starting position of pressure rise moves downstream to the end of diffuser.

Flavor perception of wine is unchanged during commercial flight: a comparative field study

Background: It is generally accepted that the human perception of flavor and odor is altered in low-pressure environments such as airplane cabins. This has been demonstrated in several simulation studies, but never in a field study conducted in an authentic environment, and never using wine as the object of study. Methods: We performed a comparative field study composed of two wine tastings. The first tasting was conducted on board an aircraft flying at standard cruising altitude and the second tasting was conducted at ground level. Subjective taste experience and current mood were evaluated through a validated questionnaire. The study was reported according to the Strengthening the Reporting of Observational Studies in Epidemiology (STROBE) guideline. Results: The study included 22 participants, none of whom had any special training in wine tasting. No statistically significant difference in experienced flavor intensity was found between the high and low altitude tests, with median aromatic intensities of 5 (interquartile range 3.5-6.0) and 5 (interquartile range 4.0-6.5) respectively, measured on a 9-point hedonic scale. Additionally, there was no detectable difference in several other taste parameters. Conclusions: These findings suggest that even though experimental studies have demonstrated that senses of taste and smell may be suppressed on commercial flights, the subjective wine tasting experience of non-professionals in real life testing may not be affected.

Famous trumpet effects

AnnotationAn analytical review of the effects obtained in the course of studies of flows in pipes is presented. There are descriptions of the famous experiments of Reynolds, who put on the agenda the problem of turbulence. The description and analysis of the discovery of Taylor vortices, which gave the received stimulus to the development of spatial hydro-gas dynamics, is given. The original invention of the Shebeko gas-dynamic tube, which allowed solving the problem of high-altitude testing of the Burya rocket, is presented. The original experiments with the thermal acoustic effect of Riike and Bossch — Riess are described, which stimulate new views on the stability of processes in the future. The invention of Gershman, which gave a practical impetus to the use of resonators in technology, is described. The description of the Rank — Hilsh pipe is given, the effect of which is widely used in the chemical and oil industry. The effect of a heat pipe used in industry for space heating is described.