Static Firing Tests Research Articles

Flow Inside a Solid Rocket Motor with Relation to Nozzle Inlet Ablation

Three-dimensional, single-phase (equilibrium two-phase) flows inside a solid rocket motor at three burn-back grain configurations are studied by computational fluid dynamics analyses of the Reynolds-averaged Navier-Stokes equations. The major concern is the relationship between the flowfield and the circumferentially periodic erosion pattern arising in the inlet region of the nozzle, which will be of help for better understanding of the surface recession mechanism. Obtained results for the first two cases show that, because the mass flux of the slot phase is notably large compared with that of the fin phase, a remarkable interphase gap in the amount of convection heating appears either in the throat or the exit cone. The peak heating rate appears, commonly to all cases, azimuthally in the slot phase and axially at the expansion ratio of about 0.9 upstream of the throat. The flow which comes out of a slot into a fin base region spreads toward the fin central part under the influence of the pressure gradient in the circumferential direction and forms a vortical flow tube of opposite rotation mutually with the flow which swirls out of the next slot At the fin phase, because the proportionality relation is accepted between the total mass recession per unit area and the total convective heat mass transfer per unit area, there is little mechanical erosion, and corrosion is considered to be dominant On the other hand, in the slot phase, surface recession which cannot be explained only by corrosion in a nozzle inlet nose exists. This surface recession has a very high possibility of having occurred by abrasion by the aluminum/alumina particles contained in the flow which comes out of the axial slot of grain and collides with the thermal protection system surface. It is expected that the periodic erosion pattern which synchronized with axial slots observed after the static-firing test is the result of such a mechanism ruling. In both the throat and the exit cone, it is thought irrespective of a phase that the effect of mechanical erosion is very small and corrosion or a so-called chemical attack is the dominant mechanism of surface recession.

Thrust-Vector Characteristics of Jet Vanes Arranged in X-Formation Within a Shroud

Thrust-vector characteristics of jet vanes arranged in an X formation within a thrust-vector-control (TVC) shroud are very unique and much more complicated than those of the jet vanes acting without the shroud because of both the flow interference between the vanes and the shroud and the X formation of the vanes. The fluid dynamic interferences, such as the impingement of shock waves induced by jet vanes on the shroud wall, and the aerodynamic performance of jet vanes arranged in an X formation were investigated to characterize thrust vector control of jet vanes operating within the TVC shroud. An integrated method consisting of a semi-empirical model, a three-dimensional numerical analysis, and static firing tests of full-scale motors was applied to quantitatively present the side-force coefficient. Satisfactory agreement is found when compared to static firing test results.

Instabilities and pressure oscillations in solid rocket motors

The purpose of this paper is to give an overview of the main results obtained on instabilities and pressure oscillations in segmented solid rocket motors. A major part of this work was carried out in the framework of the ASSM and POP R&T programs supported by the French national space agency CNES during the last decade. ASSM is related to Aerodynamics of Segmented Solid Motors and POP for Pressure Oscillations Program for the Ariane 5 solid booster (P230). Due to the use of segmented technology for the P230 motor and the possible acoustic oscillations inside the motor chamber, anticipated at the beginning of the programs and confirmed later on static firing tests, the main scientific objective of the ASSM program was oriented towards the comprehension and the modeling of the vortex shedding phenomena that are supposed to be responsible of the pressure and thrust oscillations observed in the P230. POP program was started in order to obtain an experimental and numerical data base using subscale tests of 1/15th of the P230. After the description of the instabilities observed in the P230 solid rocket booster, the scientific approach of the ASSM program is detailed insisting on the validation of numerical tools in order to predict oscillation frequencies and amplitudes. The logic of work regarding POP program is also presented. The main section of this paper provides an overview of different results obtained in ASSM and POP programs to understand the mechanisms driving to the instabilities in solid rocket chamber. The most important recent result, inside ASSM and POP programs, was the discovery of the parietal vortex shedding and the role of aluminum combustion on instabilities. Together, these two mechanisms seem to be an important potential source of instabilities and provide a new vision of the P230 stability.

3333 噴流衝突型高推力ハイブリッドロケットの開発および打上げ実験

A joint research team of universities has been organized to develop small-scale reusable launch systems based on new type of hybrid rockets. This paper describes the project outline, development study of a new type of hybrid rocket engine, and the result of a ballistic test launch using this engine. The key design of the new type hybrid rocket engine, designated as Jet Impinging type Hybrid Rocket, is that the gas flow collides with the solid fuel surface to accelerate the heat transfer to the fuel, resulting in improved thrust level. Static firing tests with an engineering model engine with a LOX cooling system showed sufficiently prompt ignition and stable combustion. Based on these results, a flight model engine was developed to conduct a ballistic test launch. The engine worked excellently in the test launch and the result was successful, showing that te engine showed expected performance in the flight condition.

Fire Resistance of Lightweight Timber Framed Walls

Limiting the fire spread through lightweight framed construction is a well-established method of providing fire safety in building design. The design objective is for walls to provide fire resistance for a certain period to satisfy fire safety requirements. This paper describes a model for predicting the fire resistance of a non-loadbearing or loadbearing timber framed wall subjected to either a standard fire resistance test or realistic fire conditions. Prediction of the fire resistance uses finite difference techniques for heat conduction within linings and for convection and radiation on the boundaries and cavity. An algorithm for the charring of timber is included to determine a structural failure condition for the studs.

720 衝突噴流式ハイブリッドロケットのフライトモデル設計のための地上燃焼試験(OS7-6 航空宇宙推進(流体,燃焼))(OS7 北海道における航空宇宙工学研究の進展と未来)

720 衝突噴流式ハイブリッドロケットのフライトモデル設計のための地上燃焼試験(OS7-6 航空宇宙推進(流体,燃焼))(OS7 北海道における航空宇宙工学研究の進展と未来)

DEVELOPMENT STUDY ON ATREX ENGINE

The study on ATREX engine (Air-Turbo Ramjet engine) development is being been conducted in ISAS since 1986 as a candidate for the propulsion system of the fly-back booster up to Mach 6 on the reusable TSTO space plane. ATREX is a fan-boosted ramjet engine using liquid hydrogen as a fuel and coolant. Sea-level static firing tests of ATREX with precooling were carried out in parallel with the wind tunnel tests on the aerodynamic components such as air intake and plug nozzle. Further studies on precooler have been conducted not only for performance improvement but also for weight reduction to meet flight requirements. In 1998, a new type precooler (Type-III) was designed and tested. Its heat transfer performance could be improved by increasing its compactness using tubes of smaller in diameter as well as its weight could be reduced. On the other hand, some difficulties increased in its manufacturing due to larger number of tubes required. The new precooler performance on heat transfer and pressure loss compared with the old type of precoolers are presented here. The test results of ATREX engine installed with new precooler is presented. Some progress in aerodynamic studies on ATREX engine is also presented here. The control of the axisymmetric air inlet by capturing the terminal shock waves and the reduction on boat tail drag of the plug nozzle were conducted in the wind tunnel.

Smoke gas analysis by Fourier transform infrared spectroscopy - summary of the SAFIR project results

The determination of toxic components from fire gases is difficult because the environment is hot, reactions are often temperature dependent, and a lot of soot may be produced. Due to the different properties of the gas components, a different time-consuming procedure for each species has traditionally been used. The use of FTIR (Fourier transform infrared) spectrometers as a continuous monitoring technique overcomes many of the problems in smoke gas analyses. FTIR offers an opportunity to set up a calibration and prediction method for each gas showing a characteristic spectral band in the infrared region of the spectrum. The objective of the SAFIR project was to further develop the FTIR gas analysis of smoke gases to be an applicable and reliable method for the determination of toxic components in combustion gases related to fire test conditions. The optimum probe design, filter parameters and the most suitable sampling lines in terms of flow rate, diameter, construction material and operating temperature have been specified. In the large scale, special concern was given to the probe design and the effects of the probe location as well as practical considerations of the sampling line length. Quantitative calibration and prediction methods have been constructed for different components present in smoke gases. Recommendations on how to deal with interferents, non-linearities and outliers have been provided and a verification method for the spectrometer for unexpected variations and for the different models have been described. FTIR measurement procedures in different fire test scenarios have been studied using the recommendations of this project for measurement techniques and analysis and an interlaboratory trial of the FTIR technique in smoke gas analysis was carried out to define the repeatability and reproducibility of the method in connection with a small scale fire test method, the cone calorimeter. Copyright © 2000 John Wiley & Sons Ltd.

Combustion behaviour of ethylene vinyl acetate copolymer‐based intumescent formulations using oxygen consumption calorimetry

This work compares the fire degradation of ethylene-vinyl acetate 8% copolymer (EVA8) with two flame-retarded formulations, using the cone calorimeter. The first one, EVA8/ammonium polyphosphate (APP) leads to the protection of the material, in the conditions of a fire, by means of blowing up and weak carbonization. The addition of polyamide-6 (PA-6) in EVA8/APP improves the protection by forming an intumescent carbonaceous shield. The fire hazard of the virgin polymer and of the FR systems are quantified, in terms of rate of heat release, weight loss, effective heat of combustion, volume of smoke production and CO and CO2 production. It assesses the effectiveness of the fire retardant additives APP and APP/PA-6 in EVA8-based materials under simulated real fire test conditions. The respective temperatures of the degradation front are deduced from the weight loss data recorded in the course of the cone calorimeter experiments and the results of the invariant kinetic parameters method applied to the different specimens. The addition of APP/PA-6 in EVA8 leads to a superficial phenomenon; the degradation zone is thus located on the upper volume of the specimen during all the exposure to the heat flux. © 1998 John Wiley & Sons, Ltd.

Combustion behaviour of ethylene vinyl acetate copolymer-based intumescent formulations using oxygen consumption calorimetry

This work compares the fire degradation of ethylene-vinyl acetate 8% copolymer (EVA8) with two flame-retarded formulations, using the cone calorimeter. The first one, EVA8/ammonium polyphosphate (APP) leads to the protection of the material, in the conditions of a fire, by means of blowing up and weak carbonization. The addition of polyamide-6 (PA-6) in EVA8/APP improves the protection by forming an intumescent carbonaceous shield. The fire hazard of the virgin polymer and of the FR systems are quantified, in terms of rate of heat release, weight loss, effective heat of combustion, volume of smoke production and CO and CO2 production. It assesses the effectiveness of the fire retardant additives APP and APP/PA-6 in EVA8-based materials under simulated real fire test conditions. The respective temperatures of the degradation front are deduced from the weight loss data recorded in the course of the cone calorimeter experiments and the results of the invariant kinetic parameters method applied to the different specimens. The addition of APP/PA-6 in EVA8 leads to a superficial phenomenon; the degradation zone is thus located on the upper volume of the specimen during all the exposure to the heat flux. © 1998 John Wiley & Sons, Ltd.

Vortex Shedding in a Large Solid Rocket Motor Without Inhibitors at the Segment Interfaces

The internal acoustic response of the Titan IV Solid Rocket Motor Upgrade (SRMU) is analyzed using pressure oscillation time histories measured during four static e ring tests. Pressure oscillations for other large solid rocket motors are caused by vortex shedding at annular inhibitors and acoustic feedback resulting from impingement of the vortices on other inhibitors or on the solid rocket motor nozzle. The SRMU does not have inhibitors, but it is shown that acoustic feedback also dee nes its pressure oscillations. Vortices are shed around the cavity between the center and aft segments and impinge on the nozzle entrance. The frequencies of the pressure oscillations vary about that of the motor fundamental acoustic mode and generally agree with a simple empirical relationship that has been used to model acoustic feedback. The effect of the SRMU pressure oscillations on dynamics and control of the Titan IV system is minor. I. Introduction A NEW solid rocket motor design called the Solid Rocket Motor Upgrade (SRMU) will be integrated with the Titan IV system beginning in late 1996. The SRMU has a lower inert weight and higher propellant weight than the current Titan IV solid rocket motor, and therefore, augments system performance. The SRMU has three segments (forward, center, and aft) and is the only large solid rocket motor in production that does not have annular inhibitors at the segment interfaces to prevent combustion and to support the propellant grain during burning. A. SRMU Static Firing Tests Five static e ring tests were conducted for qualie cation of the SRMU. The test articles included one preliminary qualie cation motor (called PQM19) and four qualie cation motors (called QM1, QM2, QM3, and QM4 ). The tests were conducted vertically, nozzle down at the Phillips Laboratory 1-125 Firing Complex, Edwards Air Force Base. 1 The test cone guration and the attachments of the solid rocket motor to the test stand simulated e ight. The physical differences between the e ve test motors were minor and are believed to have had a negligible effect on combustion stability. The static e ring tests were conducted at propellant bulk temperatures in the range 36.5‐ 937F (2.5‐ 33.97C) to assess the effect on motor performance. The static e ring test instrumentation included Condec gauges (CCC-406) of absolute pressure on the forward closure (measurement PCF1 ) and on the aft dome (measurement PCA1), and a Kistler gauge (202M122) of oscillatory pressure (measurement PD4 ) on the forward closure. 2 Measurements PCF1 and PCA1 were recorded by a Neff digital acquisition system, while measurement PD4 was recorded by a Metraplex FM system. The Neff system applied a 100-Hz low-pass e lter and recorded measurements PCF1 and PCA1 at 250 samples/s.

The development of space solid rocket motors in China

China has undertaken to research and develop composite solid propellant rocket motors since 1958. At the request of the development of space technology, composite solid propellant rocket motor has developed from small to large, step by step. For the past thirty eight years, much progress has made, many technical obstacles, such as motor design, case materials and their processing technology, propellant formulations and manufacture, nozzles and thrust vector control, safe ignition, environment tests, nondestructive inspection and quality assurance, static firing test and measurement etc. have been solved. A serial of solid rocket motors have been offered for China's satellites launch. The systems of research, design, test and manufacture of solid rocket motors have been formed.

A Comparative Study of the Fire Performance of Halogenated and Non-Halogenated Materials for Cable Applications. Part II Tests on Cable

A total of 21 electrical cables were made, all with essentially identical construction but differing in the chemical composition of sheath and/or insulation, which were all commercially available materials, both halogenated and non-halogenated. All cables were tested in two large-scale cable tray tests, ASTM D5424 (CSA FT-4 protocol), with a total length of 2.44 m and IEC 332-3, with a total length of 3.5 m. The cables were also tested in a number of small- and medium-scale tests for flame spread (IEC 695-2-2, IEC 332-1, UL 1581 Part VW1, BS 476 Part 12E, DIN 4102 Part 16), temperature increase (DIN 4102 Part 16) and smoke obscuration (IEC 1034-2, BS 476 Part 12E). Finally, all cables were tested in the cone calorimeter (ISO 5660), horizontally, at incident fluxes of 20, 40 and 70 kW m−2. All the cables passed the mild flammability tests, but distinctions could be made based on the afterflame time observed, where halogenated cables outperformed non-halogenated cables by a significant margin. It was also possible to distinguish between the halogenated and non-halogenated cables on the basis of the cable length charred in some tests. In terms of smoke obscuration, it was found that the rankings offered by the various tests were very different. While non-halogenated cables had improved smoke performance over traditional vinyl types, fluorinated cables performed very well. This confirms the importance of material selection by performance rather than by chemical composition. Almost all cables performed sufficiently well that they generated relatively limited amounts of smoke under realistic end-use fire test conditions. The peak heat release rate in the large-scale cable tray test (ASTM D5424) served as an excellent criterion for discriminating between the fire performance of the various cables (the traditional criterion being char length). The average rate of heat released also served to distinguish between different levels of cable fire performance. Moreover, cables passing the test tended to release less heat and smoke than those that failed. The trends observed in the cone calorimeter heat release test were similar to those in the large-scale test and show good correlation between cable tray char length and cone calorimeter heat release. It was observed that the halogenated cables tested performed better than the non-halogenated cables in terms of heat release rate by factors ranging from two to greater than five. The results indicate that cables with excellent fire performance can be constructed by using a variety of materials. It is thus important to specify fire performance and leave material choice to manufacturers.

Space Shuttle redesigned solid rocket booster structural dynamic liftoff predictions and correlations

A transient analysis was performed that predicted and correlated the structural static and dynamic behavior of the Space Shuttle redesigned solid-rocket booster (RSRB) during the first redesigned flight, STS-26R, liftoff. A finite element model of the RSRB was analyzed to predict the accelerations, strains, and displacements measured by developmental flight instrumentation that was bonded to the motor during flight. A detailed RSRB finite element model was developed and correlated for use in analyzing the booster behavior during liftoff loading conditions. This finite element model used MSC/NASTRAN superelement techniques to model all components of the RSRB in detail. The accuracy of the assembled RSRB global model was validated by correlation to static fire and other tests performed by Thiokol Corporation. Mode acceleration transient analysis methods were applied to this validated RSRB global model to predict RSRB structural behavior using the actual loading conditions measured during liftoff. Essential information describing the analytical model and techniques used, correlation of the predicted vs measured RSRB behavior, and conclusion are presented in this study.

Evaluation of kinetic parameters appropriate for modeling urethane foam insulation performance

Computer codes which model the pyrolysis of thermal insulators for in-depth temperature response are particularly sensitive to the kinetic parameters used in the code. The parameter values, which are evaluated from thermogravimetric analysis (TG), are, in turn, sensitive to the heating and gas flow rates in the TG sample environment as well as the composition of the gas. In this study, the effects of these experimental variables on computer code predictions and their correlations with experimental data are demonstrated. The insulators are urethane foams and were subjected to simulated and actual instantaneous fire test conditions. The Charring Material Ablation computer code was used to model insulator performance.

Experiments with solid rocket technology in the development of M-3SII

Covered topics concern with significant problems for vehicle's performance of M-3SII which has been developed by ISAS to send interplanetary probes towards Halley's comet. Characteristic features of each motor are described with design principles of subsystems as well as experiments in static firing tests. The importance of gas-particle interaction effects is emphasized in some respects as residual thrust, erosion of insulation liner, safety for debris dispersion and so forth.

Computation of the Shuttle solid booster nozzle start-up transient flow

The first NASA Space Shuttle flight (STS-1) produced an overpressure wave that exceeded preflight predictions by as much as 5 to 1. This second overpressure wave occurred just after the first overpressure wave produced by the lolid rocket booster igniter. To understand this second overpressure phenomenon, a numerical simulation effort was undertaken. Both the SRB static firing test and STS-1 geometries were studied for twodimensional (axisymmetric), inviscid and viscous flow. The inviscid calculations did not produce significant second overpressure waves. However, the viscous calculations did produce second overpressure waves that qualitatively agree with experiment. These overpressure waves were present in both the static firing test and STS-1 geometries. This second overpressure wave is generated by the motion of the boundary-layer separation point and the subsequent radial motion of the exhaust jet during the start-up of the SRB nozzle flow. The presence of the mobile launch platform exhaust hole wall amplifies this wave, but does not appear to be the source of any additional overpressure waves. The lack of good quantitative agreement between theory and experiment indicates that other overpressure sources, not accounted for by this simulation, may be present.

ロケット・エンジン真空燃焼実験

ロケット・エンジン真空燃焼実験

Electronic Devices for Hydrogen Leak and Fire Detection

Electronic hydrogen leak and fire detector systems are installed on the Saturn S-IVB static test firing stands to detect hydrogen leaks and hydrogen fires. Only this portion of the hydrogen safety program is covered in this paper. The characteristics of hydrogen and hydrogen flames are reviewed. The results of the investigations into various types of leak and fire detectors is recorded. Installation and operational problems and the solutions to these problems are discussed.

Air pollution evaluation of Titan II test firings.

Abstract An air pollution evaluation was made at the site of Titan II test firings to determine the downwind personnel hazard from exhaust gases. Hundreds of thousands of pounds of propellants are used in a test firing. Until sampling results proved otherwise, the pollution potential from a firing appeared tremendous. A study of the toxicologically significant contaminants of the environment near test-stand facilities for the Titan II missile engines was conducted to determine the downwind concentrations during static-test firing.