Hot-fire Tests Research Articles

Acoustic signature characterization of a sub-scale rocket launch

Georgia Tech Research Institute (GTRI) conducted a flight test of a sub-scale rocket in 2013 outside Talladega, Alabama, to acquire the launch acoustics produced. The primary objective of the test was to characterize the acquired data during a sub-scale launch and compare it with heritage launch data from the STS-1 Space Shuttle flight. Neither launch included acoustic suppression; however, there were differences in the ground geometry. STS-1 launched from the Mobile Launch Platform at Pad 39B with the RS-25 liquid engines and Solid Rocket Boosters (SRBs) firing into their respective exhaust ducts and flame trench, while the GTRI flight test vehicle launched from a flat reflective surface. The GTRI launch vehicle used a properly scaled Solid Rocket Motor (SRM) for propellant; therefore, primary analysis will focus on SRM/SRB centric acoustic events. Differences in the Ignition Overpressure (IOP) wave signature between both due to this will be addressed. Additionally, the classic liftoff acoustics “football shape” is preserved between both full and sub-scale flights. The launch signatures will be compared, with note taken of specific launch acoustic events more easily investigated with sub-scale launch data or supplement current sub-scale static hotfire testing.

Combustion Characteristics of a Coaxial Porous Injector

A coaxial porous injector was proposed for gas–liquid combustion for improved mixing performance compared to a conventional shear coaxial injector. The coaxial porous injector discharged a gas jet in a direction normal to the center liquid core, and it was surmised that the momentum transfer from the radial gas jet to the axial liquid jet occurred more efficiently than with a traditional shear coaxial injector. To validate the effectiveness of this idea, the combustion performances of a coaxial porous injector and a shear coaxial injector were compared by conducting a hot-firing test using nitrous oxide-ethanol as propellant. The momentum flux ratio at the injector tip was selected as a spraying condition index, and the characteristic velocity efficiency was considered as a combustion performance parameter. The coaxial porous injector demonstrated higher combustion efficiency than the conventional shear coaxial injector, especially at lower-momentum flux ratio conditions. In addition, the comparison of spray images between the shear coaxial and coaxial porous injectors under ambient conditions supported the results of the hot-firing experiment.

Development of an Aeroelastic Modeling Capability for Transient Nozzle Flow Analysis

Lateral nozzle forces are known to cause severe structural damage to any new rocket engine in development during testing. Although three-dimensional, transient, turbulent, chemically reacting computational fluid dynamics methodology has been demonstrated to capture major side load physics with rigid nozzles, hot-fire tests often show nozzle structure deformation during major side load events, leading to structural damages if structural strengthening measures were not taken. The modeling picture is incomplete without the capability to address the two-way responses between the structure and fluid. The objective of this study is to develop a coupled aeroelastic modeling capability by implementing the necessary structural dynamics component into an anchored computational fluid dynamics methodology. The computational fluid dynamics component is based on an unstructured-grid pressure-based computational fluid dynamics formulation, whereas the computational structural dynamics component is developed under the framework of modal analysis. Transient aeroelastic nozzle startup analyses at sea level were performed to demonstrate the successful simulation of nozzle wall deformation with the proposed tightly coupled algorithm, and the computed results pertinent to fluid–structure interaction presented.

연소불안정에 따른 축소형 연소기에서의 열전달 영향

동축 와류형 분사기 19개로 구성된 연소기 헤드와 냉각채널을 갖는 연소실을 이용하여 연소시험을 수행하였다. 추진제로는 액체산소와 케로신(Jet A-1)이 사용되었으며, 연소시험은 연소실 압력 59~82 bar, 혼합비 2.0~3.0 영역에서 수행되었다. 냉각채널 연소실의 냉각 유체로는 물이 사용되었으며, 냉각채널 입구와 출구에서의 물의 온도를 측정하여 열유속 값을 계산하였다. 본 연구에서는 연소불안정에 따른 열전달 영향을 살펴보는 것을 목표로 하였으며, 이를 위해 냉각수의 온도 변화를 계측하였다. 몇 번의 연소시험에서 연소불안정 현상이 발생하였으며, 이때 열유속이 5~20% 정도 증가하는 결과가 나타났다. 또한 열유속은 연소불안정이 발생하는 초기 시점에서 최대가 되는 것을 알 수 있었다. Hot-firing tests were carried out using a mixing head with 19 swirl coaxial injectors and a combustion chamber with internal cooling channels. The propellants of liquid oxygen and kerosene(Jet A-1) were burned in a range of chamber pressures (59~82 bar) and mixture ratios (2.0~3.0). The temperature of water used as the cooling fluid was measured at the inlet and outlet of the cooling channels, and the heat flux was calculated. The aim of this study was to examine the effect of combustion instability on heat transfer in a subscale thrust chamber, and detect the temperature variation of cooling water. During several hot-firing tests, combustion instability was encountered which caused a 5~20% increase in heat flux. The peak heat flux took place in the initial stages of combustion instability.

Effective Modeling of Conjugate Heat Transfer and Hydraulics for the Regenerative Cooling Design of Kerosene Rocket Engines

The present work is motivated to develop a unified framework to simulate multi-physical processes which are crucial for trade-off design of liquid rocket thrust chambers among propulsive performance, regenerative cooling, and pressure budget. In this paper, an effective modeling of conjugate heat transfer and hydraulics through the regenerative cooling passage has been performed to quantitatively evaluate detailed cooling designs, including spirally twisted channels and bidirectionally branched circuit, as well as to provide the wall heat flux to a compressible reacting flow solver in an interactively coupled manner. The kerosene fuel used as coolant is modeled by a three-component physical surrogate, and the fluid properties required for calculation of a Nusselt number correlation and empirical resistance coefficients are computed over the entire thermodynamic states from compressed liquid to supercritical fluid using the NIST SUPERTRAPP. The present method has been applied to an actual regeneratively cooled thrust chamber and validated against measurement of hot-firing tests in terms of temperature increase and pressure drop of the coolant through the cooling passages. Based on the numerical results, supplementary effects of peripheral fuel cooling injection and thermal barrier coating are addressed.

액체로켓엔진용 파이로시동기의 설계 및 연소시험연구

액체로켓엔진용 파이로시동기의 설계 및 연소시험연구

Coupling of Cryogenic Oxygen–Hydrogen Flames to Longitudinal and Transverse Acoustic Instabilities

A rectangular combustor with acoustic forcing was used to study flame–acoustic interaction under injection conditions that are representative of liquid propellant rocket engines. Hot-fire tests using liquid oxygen and gaseous hydrogen were conducted at pressures of 40 and 60 bar, which are sub- and supercritical conditions, respectively, for oxygen. Examined samples of hydroxyl-radical emission imaging, collected using a high-speed camera during periods of forced acoustic resonance, showed significant response in the multi-injection element flame. Fluctuating acoustic pressure causes in-phase fluctuation of the emission intensity, producing response factor values of around 0.8. Transverse acoustic velocity causes shortening of the flame, concentrating heat release near the injection plane. The flame is also convectively displaced with transverse acoustic velocity, a process believed by many to be responsible for driving transverse mode high-frequency combustion instabilities. The analysis in this work was extended to detect the response of internal flame processes to the transverse acoustic disturbance, namely, pressure response and enhanced transport and mixing of non-premixed propellants. A flame-tracking technique was implemented to isolate the response of these processes from that of collective convective displacement. The results show a significant contribution to overall flame response from the response of internal processes; however, further work is required to quantify the relative contributions of pressure or propellant transport.

액체로켓엔진 연소기 연소안정성 평가시험

액체로켓엔진 연소기의 연소안정성을 평가하기 위한 방법으로 연소 시 외부 교란을 공급하여 안정성 특성을 확인하는 방법이 사용된다. 본 논문에서는 연소기의 연소안정성 평가시험을 위한 펄스건 기폭시험에 대한 연구를 수행하였다. 비연소조건에서 펄스건 기폭시험을 수행하여 압력파의 강도를 확인함으로써 펄스건의 장약량을 결정하였다. 동일한 연소기 헤드에서 펄스건이 적용되지 않은 14번의 연소시험 동안 연소불안정이 나타나지 않았지만, 펄스건이 적용된 두 번의 탈설계점 조건에서는 연소불안정 현상이 발생하였다. 이로부터 펄스건을 이용한 기폭시험이 연소안정성 평가의 방법이 될 수 있음을 확인하였으며, 연소안정성 판정을 위한 연소시험 횟수를 줄일 수 있는 방법임을 제시하였다. As a evaluation method of combustion stability in a liquid rocket engine thrust chamber, external disturbance devices are used. In the paper, the study on pulse-gun ignition tests for a combustion stability rating test of a thrust chamber was performed. Charging volume of pulse-guns was determined by confirming the intensities of the pressure waves from the ignition tests in the cold-flow conditions. While using same injector head, combustion instabilities were not encountered during 14 hot-firing tests without pulse-guns but combustion instabilities were triggered by pulse-gun ignition during 2 hot-firing tests. The results showed that the pulse-gun ignition test could be the evaluation method and could reduce the hot-firing test number for the stability rating of a thrust chamber.

액체로켓 엔진 연소기의 열차폐 코팅 및 막냉각 조건에 따른 냉각 성능 변화 해석

The effect of ceramic thermal barrier coating thickness on the cooling performance of a liquid-propellant rocket engine combustor has been investigated through combustion/cooling performance analysis whose results verified against measured data from hot-firing tests. Also have been confirmed the effects of film cooling amount near the face plate on the coolant temperature and on the thermal barrier coating surface temperature. Some important points to be considered for designing cooling schemes for regeneratively cooled rocket engine combustor have been drawn and reviewed from present study and further verification of the analysis tool should be performed in the future.

Degradation of CVD-SiC coated PIP-C/SiC composites exposed to the hot firing test based on MMH/N2O4 bipropellants

Ablation property of CVD-SiC coated PIP-C/SiC composites was investigated by using a hot-firing test in a liquid propellant rocket motor. The effect of the combustion temperature was studied; then ablation mechanisms of coated composites were discussed. The degradation of PIP-C/SiC composites was determined by analyzing surface variations. Two different ablation modes of C/SiC composites were discussed. The coupling effects (thermal–mechanical denudation and thermal–chemical erosion) worsened the degradation of the coated PIP-C/SiC composites. CVD-SiC coating provided better protection for the composites under the MMH/N2O4 bipropellants tests.

Coupled Heat Transfer Analysis in Regeneratively Cooled Thrust Chambers

The coupled hot gas–wall–coolant environment that occurs in regeneratively cooled liquid rocket engines is studied by a computational procedure able to provide a quick and reliable prediction of thrust chamber wall temperature and heat flux as well as coolant flow characteristics, like pressure drop and temperature gain in the regenerative circuit. The coupled analysis is performed by means of a computational fluid dynamics solver of the Reynolds-averaged Navier–Stokes equations for the hot gas flow and by a simplified quasi-two-dimensional approach, which widely relies on semiempirical relations, for the coolant flow and wall structure heat transfer in the cooling channels. Coupled computations of the space shuttle main engine main combustion chamber are performed and compared with available literature data. Results show a reasonable agreement in terms of coolant pressure drop and temperature gain with nominal data, whereas the computed wall temperature peak is closer to hot-firing test data than to the nominal value.

70 N급 하이드라진 추력기의 촉매대 형상(L/D) 최적화 연구

우주비행체의 주 엔진 및 자세제어용 추력기로 적용 가능한 단일액체추진제 하이드라진 추력기의 촉매대 최적설계형상 도출을 위한 지상연소시험을 수행하였다. 추력, 비추력, 연소실 압력, 특성속도, 그리고 하이드라진 분해반응률 등을 추력실 길이변화에 따라 고찰하고, 추진제 공급압력과 성능변수가 갖는 상관관계를 제시한다. 성능평가결과, 시험조건 내에서의 촉매대 길이증가는 추력기의 성능저하를 야기하고, 공급압력 변이에 따른 추진제 소모 효율성 또한 떨어뜨리는 것이 확인되었다. A ground hot-firing test was conducted to take out the optimal design configurations for the catalyst bed of liquid-monopropellant hydrazine thruster which could be used for primary engine or attitude control thruster of space vehicles. Performance characteristics with the variation of thrust-chamber length are investigated in terms of thrust, specific impulse, chamber pressure, characteristic velocity, and hydrazine decomposition rate. Additionally, the correlations between propellant-supply pressure and performance parameters are given. As results, increase of catalyst-bed length leads to performance degradation in this test condition, and also decreases propellant consumption efficiency with the supply pressure variation.

Multidisciplinary simulation of a regeneratively cooled thrust chamber of liquid rocket engine: Turbulent combustion and nozzle flow

The present study proposed a unified framework to simulate multi-physical processes which are crucial for trade-off design of liquid rocket thrust chambers among propulsive performance, regenerative cooling, and pressure budget. As the first part, a turbulent combustion model based on the flamelet approach was developed to effectively incorporate detailed chemistry of high hydrocarbon fuel, turbulent mixing, enthalpy loss, and pressure variations within the nonadiabatic nozzle flow. In order to correctly capture the convective heat transfer and viscous friction in the turbulent boundary layer at the chamber wall, an advanced low-Reynolds number turbulence model is adopted in an axisymmetric compressible RANS solver, which is interactively coupled with a cooling analysis module for the conjugate heat transfer and hydraulics through the regenerative cooling channels. The present method has been applied to an actual regeneratively cooled thrust chamber and compared with measurement of hot-firing tests in terms of specific impulse, characteristic velocity, and thrust coefficient. Based on the numerical results, the effects of additional fuel cooling injection and wall friction on the propulsive parameters are discussed.

동축형 다공성재 분사기의 에탄올/아산화질소 연소성능

The gas jet from a coaxial porous injector for two-phase flows is discharged radially from the porous surface, which encloses the center liquid jet. Several hot-firing test using ethanol/nitrous oxide propellants was conducted to analyze the effect of oxidizer/fuel ratio on the combustion performance, and the uncertainty analysis was performed for the results. The characteristic velocity was affected by oxidizer/fuel ratio similarly with the results of CEA calculation except that the maximum characteristic velocity was appeared in the stoichiometric ratio. The characteristic velocity efficiency was increased as the oxidizer/fuel ratio increases.

Erosion of carbon/carbon composites using a low-velocity, high-particle-concentration two-phase jet in a solid rocket motor

The erosion of a four-direction carbon/carbon composite test piece using a low-velocity, high-particle-concentration two-phase jet was studied by the hot firing test of a small solid rocket motor with an elaborately designed flow path. The linear ablation rates were measured. The ablation surface and microstructure of the carbon/carbon composites were studied by scanning electron microscopy, and the ablation mechanism was investigated. Within the parameter range studied in this paper, mechanical erosion is found to play a dominant role in the ablation of carbon/carbon composites in the impact region. Moreover, the effect of chemical ablation is weak. The erosion of carbon/carbon composites results in blunt fracture tip fibers and a lamellar matrix. The particle impact mass flux is more dominant than the particle impact velocity in the erosion of carbon/carbon composites. At mesoscale, the carbon rods are more resistant to mechanical erosion than fiber bundles. At microscale, the carbon fibers are more susceptible to mechanical erosion than the carbon matrix, whereas the carbon fibers are more resistant to chemical ablation than the matrix and interface.

Probabilistic Methods to Determine Resonance Risk and Damping for Rocket Turbine Blades

The liquid oxygen–hydrogen J-2X rocket engine, which powers the upper stage of the NASA Space Launch System, is presently undergoing hot-fire testing. Analysis revealed numerous potential resonance issues with hardware in the turbomachinery flowpath, which could be catastrophic for the launch vehicle. This paper describes a series of probabilistic analyses incorporating nondeterministic characteristics of mistuning amplification, blade damping, natural frequency, and speed to assess the risk of failure of turbine blades due to resonance during past and present tests and to determine an appropriate value of damping to use in deterministic forced response analysis. The analyses are the first reported in the literature combining these critical nondeterministic parameters. The new damping determination technique developed here indicates that combining a damping value in the lower 1.5 sigma range with a 3 sigma mistuning amplification value results in a desired total 3 sigma probability of failure. These risk calculations were invaluable for determining high-cycle-fatigue margins of safety and in determining if risk-reduction methods such as dampers were necessary immediately, or if tests could be performed before risk-reduction hardware was ready.

Thermites for Solid Rocket Motor Ignition

Fast and repeatable ignition transients for small solid rocket combustors can be difficult to achieve. This work sets out to characterize a thermite mixture to fill this need. Igniter function was determined using high-speed imaging, allowing an examination of product droplet size and combustion time as a function of packaging technique. Safety testing (electrostatic discharge, drop weight impact, and friction) indicated that this material is far safer than existing ignition compounds. A simple dual-criteria ignition model is applied to igniter sizing, and our modeling results were successfully evaluated using hot fire tests of motors with various exposed propellant surface conditions. The end result is a safe, inexpensive, reliable, and readily available method of igniting small (500 g–50 kg propellant mass) solid rocket combustors.

GS38 液体ロケットエンジン外筒付き燃焼室の破損メカニズムの解明 : 上段エンジンのき裂発生メカニズム

In developing liquid rocket engines, life time of a combustion chamber is one of key design issues. In order to understand and predict the mechanism of failure mode caused by creep-fatigue, which is one of major risks to a combustion chamber, a new chamber life prediction methodology, based on a multi-physics two-way coupled simulation and the linear cumulative damage rule, was developed. The developed methodology was applied to a regeneratively cooled thrust chamber of an actual engine, and the deformation mechanism of creep-fatigue crack occurred around the throat during hot firing tests was investigated in detail.

액체로켓엔진 재생냉각 연소기의 점화 특성 연구

국내 기술로 개발된 액체로켓엔진 연소기의 점화 특성을 연소시험을 통해 분석하였다. 연소시험에 사용된 연소기들은 점화제용 분사기, 막냉각 방식 그리고 산화제 입구 위치가 다른 특징을 지니며 각기 다른 시동 시퀀스를 적용하였다. 분석결과 산화제 매니폴드의 온도 프로파일에 따라 저주파 섭동이 점화구간 내에서 나타남을 보였다. 이 저주파 섭동은 연소기의 기능장애를 유발하지는 않았지만, 엔진 시스템 및 발사체와의 인터페이스측면에서 지속적인 관심 요소로서 관찰이 요구된다. The ignition characteristics of liquid rocket engine thrust chambers which have been developed by domestic technology were analyzed through hot-firing tests. Thrust chambers used in hot-firing tests have different characteristics in terms of the injector for ignition, film cooling method and the position of the oxidizer inlet. Also, these thrust chambers used their respective startup sequences. Analysis results showed that according to temperature profiles of the oxidizer manifold, low frequency fluctuation was appeared in ignition area. This low frequency fluctuation didn't give rise to violent malfunction of the thrust chamber, but the continuous observation as a concern parameter in the side of interfaces with engine system and launch vehicle should be demanded.

우주발사체 자세제어용 하이드라진 추력기의 정상상태 추력 특성

우주발사체 자세제어용 하이드라진 추력기의 지상연소시험을 수행하였다. 시험에 사용된 추력기는 추진제 주입압력 2.41 MPa (350 psia) 에서 정상상태 공칭추력 67 N (15 <TEX>$lb_f$</TEX>) 을 목표로 설계/제작 되었다. 개발모델 추력기의 성능특성 검토를 위해 정상상태 연소모드에서의 추력, 추진제 공급압력, 질량유량, 추력실 압력, 그리고 온도 등의 성능변수를 이용한다. 시험결과, 실제의 성능이 이론 요구규격 대비 89.1% 이상의 성능효율을 만족하는 것이 확인되었다. An ambient hot-firing test was carried out for the hydrazine thruster which may be employed in the space launch vehicles. The thruster is designed to produce 67 N (15 <TEX>$lb_f$</TEX>) of nominal steady-state thrust at an inlet pressure of 2.41 MPa (350 psia). A scrutiny into the performance characteristics of thruster is made in terms of thrust, propellant supply pressure, mass flow rate, chamber pressure, and temperature at the steady-state firing mode. As a result, it is ensured that the practical performance efficiencies are above 89.1% compared to its ideal requirements.