Simulations of High Altitude Tests for Large Area Ratio Rocket Motors
25 Jan 2013-AIAA Journal (American Institute of Aeronautics and Astronautics)-Vol. 51, Iss: 2, pp 433-443
TL;DR: In this article, the performance characteristics of a high altitude test facility for testing large area ratio rocket engines have been investigated, and the predicted results show that the desired vacuum level is attained when the primary jet flow attaches to the ejector duct walls smoothly, thereby arresting any back flow.
Abstract: In the present study, performance characteristics of a high altitude test facility for testing large area ratio rocket engines have been investigated. Steady-state numerical simulations have been performed initially to highlight the effects of operational parameters on a high altitude test facility operation. Later, the performance of the test facility during the startup phase of the rocket motor has been analyzed. The predicted results show that during the initial high altitude test facility evacuation, the desired vacuum level is attained when the primary jet flow attaches to the ejector duct walls smoothly, thereby arresting any back flow. However, at the fully started condition of the motor, the self-ejector action of the rocket plume plays a major role in maintaining the desired vacuum condition and, hence, the ejector flow rate can be reduced significantly. The injection of water as a fine spray cools the hot gas to a sufficiently low temperature (∼600 K) prior to its release into the atmosphere. T...
TL;DR: In this article, the effect of the initial expansion angle of a parabolic (TOP) nozzles on flow separation pattern and shock structure was investigated numerically, and it was shown that the presence of a restricted shock separation pattern leads to a considerable increment of the critical cross sectional area of the flow inside the diffuser, and the minimum starting pressure of the STED is increased up to 30% after the resizing of the second throat area to eliminate the flow choking inside it.
Abstract: Free or restricted shock separation phenomena can occur inside a thrust optimized parabolic (TOP) nozzle during over-expanded operations. In the case of restricted shock separation, a cap shock pattern forms in the nozzle which leads to a substantial total pressure drop. This induces further related issues in the process of ground testing of such nozzles using a second throat exhaust diffuser (STED). In the present study, the flow physics in several TOP nozzles operating at over-expanded conditions is investigated numerically. At first, the strong effect of the initial expansion angle of a TOP nozzle on flow separation pattern and shock structure is demonstrated. Results reveal that for high initial expansion angles, restricted shock separation occurs even at low nozzle pressure ratios, while free shock separation takes place for small initial expansion angles at even high nozzle pressure ratios. Subsequently, the effect of separation patterns in TOP nozzles on the starting pressure of a STED is studied. Current results show that the presence of a restricted shock separation pattern leads to a considerable increment of the critical cross sectional area of the flow inside the diffuser. Therefore, the minimum starting pressure of the STED is increased up to 30% after the resizing of the second throat area to eliminate the flow choking inside it.
TL;DR: In this article, an experimental analysis of a second-throat exhaust diffuser (STED) performance has been conducted for the high altitude test of a parabolic bell-type nozzle.
Abstract: In the present study, the experimental analysis of a second-throat exhaust diffuser (STED) performance has been conducted for the high altitude test of a parabolic bell-type nozzle. The diffuser starting performance has been explored considering two different approaches, namely the gradual and instantaneous increase of the nozzle chamber pressure. Also, the influence of pre-evacuating the internal regions of the system on the diffuser starting performance has been studied. Numerical simulations have been carried out to have more physical insight into some test results. It is demonstrated that the non-dimensional hysteresis range in STED starting performance with a bell-type nozzle is up to 4 times larger than that with a conical nozzle. The pre-evacuation of the internal area does not affect the diffuser minimum starting pressure. However, starting the diffuser with pre-evacuation takes 50 to 70 percent less time than without pre-evacuation. Also, it is illustrated that in spite of large differences in minimum starting pressure of STED with conical and bell-type nozzles, the STED breakdown pressure (minimum operating pressure) is independent of nozzle profile. Moreover, it is shown with experimental evidences and numerical analyses that the creation of restricted shock separation (RSS) inside the bell nozzle leads to some unexpected behaviors in STED starting and breakdown behaviors. Finally, it is demonstrated that the instantaneous increase in the nozzle chamber pressure with steeper slope and pre-evacuation of the internal regions can eliminate separation pattern transition from common free shock separation to unwanted RSS.
10 Nov 2018
TL;DR: It is confirmed that the World Health Organization has approved the use of nanofiltration membranes for the recovery of phosphorous-contaminated wastewater from the Mediterranean Sea and Atlantic Ocean.
Abstract: هلاقم تاعلاطا هدیکچ لماک یشهوژپ هلاقم :تفايرد 24 نمهب 1396 :شريذپ 04 نيدرورف 1397 :تياس رد هئارا 07 تشهبيدرا 1397 کي درکلمع رضاح قیقحت رد هیبش هاگشيامزآ عافترا زاس هیبش درکيور اب روتوم کي ندش شوماخ نامز رد رارق یسررب دروم نايرج یددع یزاس یلصا هتسه رد یقارتحا یاهزاگ نايرج مئادریغ یددع لیلحت .تسا هتفرگ هیبش عافترا زاس شيامزآ هظفحم و هيوناث هاگولگ رزویفيد( ) لیفورپ اب راشف رد هدافتسا دروم رزویفيد درکلمع و متسیس نيا رد نايرج کيزیف .تسا هدش ماجنا روتوم یشوماخ نامز یقارتحا یاهزاگ راکدوخ هیلخت یسررب .تسا هتفرگ رارق یسررب دروم یم ناشن رضاح یددع یاه یلع هک دنهد لااب قارتحا راشف رد روبزم رزویفيد هکنيا مغر ي ی هار یم یزادنا ،دوش نیياپ بتارم هب قارتحا راشف رد روتوم یشوماخ نامز رد هار تلاح زا یرت یم جراخ یزادنا تشگرب اب .دوش ،شيامزآ هظفحم لخاد هب مرگ یاهزاگ ات هظفحم لخاد لایس طسوتم یامد K 2200 یم شيازفا یاهرازبا هب تسا نکمم شيامزآ هظفحم لخاد لایس یامد شيازفا نازیم نيا .دباي هزادنا فلتخم عافترا اب یتشگرب نايرج هدننک دودحم بصن ریثات ،قیقحت همادا رد .دناسرب بیسآ نآ رد دوجوم یریگ لخاد لایس یيامد تارییغت رد طسوتم یامد شهاک رد عنام نيا بصن ریثات هک تسا هدش هداد ناشن .تسا هتفرگ رارق یسررب دروم روتوم یشوماخ نامز رد شيامزآ هظفحم مشچ شيامزآ هظفحم لایس یروط هب ،تسا ریگ س کي زا رتمک هب هظفحم رد لایس یامد ،بسانم عافترا اب عنام کي بصن اب هک تلاح ربارب مو یم شهاک هدننک دودحم زا هدافتسا نودب .دباي :ناگژاو دیلک هیبش هاگشيامزآ عافترا زاس توص قوفام رزویفيد روتوم یشوماخ هلحرم هیبش یددع یزاس یتشگرب نايرج هدننک دودحم عنام
01 Jan 2016
TL;DR: In this article, the influence of regional pre-evacuation of an altitude test facility on the starting time of a second throat supersonic exhaust diffuser is numerically investigated.
Abstract: In this paper, the influences of regional pre-evacuation of an altitude test facility on starting time of a second throat supersonic exhaust diffuser are numerically investigated. Detailed numerical studies have been carried out to evaluate the physics of the flow and starting time of the diffuser for different pre-evacuation zones along the diffuser and the test chamber. Unsteady axisymmetric compressible Navier–Stokes equations, incorporated with the two equation kω-SST turbulence model are solved, with density-based solver to extract the current flow features. The numerical method is properly validated with the measured data available in the literature. Our investigations show that the amount of preevacuation volume has strong effects on starting time of the diffuser. As we extend pre-evacuation zone along the diffuser, the smaller starting time of diffuser is resulted. However, the increasing of pre-evacuated test chamber size causes the increasing of starting time of the diffuser.
24 Mar 1923
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).
TL;DR: In this article, wall static and in-stream phot pressure distributions are presented for confined, nonreacting, supersonic flows in cylindrical sections wherein a shock structure has been stabilized.
Abstract: Wall static and in-stream phot pressure distributions are presented for confined, nonreacting, supersonic flows in cylindrical sections wherein a shock structure has been stabilized. Based on an analysis of these measurements, the character of the wave structure is shown to be oblique rather than normal, with the flow remaining primarily supersonic downstream of the shock system. When additional cylindrical sections are either added or deleted the shock structure is, with the exception of slight changes due to the different initial conditions, independent of location in the duct. The parameters which govern the distance st, over which the pressure rise is spread, viz., Mach number, momentum thickness Reynolds number, duct diameter, and the momentum thickness of the upstream boundary layer, were varied as follows: 1.53 ^ Ma ^ 2.72, 5 x 10 ^ Ree ^ 6 x 10, 1.0 D 6.1 in., and 0.007 ^ 6 ^ 0.036 in. In each test the wave structure was generated by either lowering the pressure in the air supply system so that the cylindrical duct was, in effect, overexpanded when discharging to ambient conditions, or by throttling the flow leaving the duct. For a given pressure ratio across the disturbance, Pf/pa, st varies approximately directly with the product 0D and inversely with (Ma — l)Re0. A simple quadratic expression is presented which adequately represents this corespondence for the complete range of conditions tested and for data from the cited reference.
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 .
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.