Rocket Combustion Chamber Research Articles

Life prediction of thermally highly loaded components: modelling the damage process of a rocket combustion chamber hot wall

During their operational life-time, actively cooled liners of cryogenic combustion chambers are known to exhibit a characteristic so-called doghouse deformation, pursued by formation of axial cracks. The present work aims at developing a model that quantitatively accounts for this failure mechanism. High-temperature material behaviour is characterised in a test programme and it is shown that stress relaxation, strain rate dependence, isotropic and kinematic hardening as well as material ageing have to be taken into account in the model formulation. From fracture surface analyses of a thrust chamber it is concluded that the failure mode of the hot wall ligament at the tip of the doghouse is related to ductile rupture. A material model is proposed that captures all stated effects. Basing on the concept of continuum damage mechanics, the model is further extended to incorporate softening effects due to material degradation. The model is assessed on experimental data and quantitative agreement is established for all tests available. A 3D finite element thermo-mechanical analysis is performed on a representative thrust chamber applying the developed material-damage model. The simulation successfully captures the observed accrued thinning of the hot wall and quantitatively reproduces the doghouse deformation.

A nanostructured ablative bulk molding compound: Development and characterization

Glass and silica fiber phenolic composite ablators are effectively used in the production of passively cooled rocket combustion chambers. To improve the performance of these composites as well as to decrease their costs, the use of micron sized silicon dioxide particles has been widely adopted. In this work, the possibility of producing glass phenolic composites with nanosized silica as an alternative to micron-scaled silicon dioxide was investigated. The ablative properties of the produced materials were studied using an oxy-acetylene torch. In depth temperature profiles taken through the thickness of the samples, loss of mass data and an optical analysis of the post-test surfaces were used to evaluate the effects of nanosilica. Furthermore, to investigate the material post-test microstructure, a detailed morphological characterization was carried out using Scanning Electron Microscopy (SEM). In comparison to neat glass/phenolic composites, the introduction of nanosized silica particles embedded in the matrix significantly improved the ablative properties in terms of mass loss and erosion rate.

Experimental research on the rotating detonation in gaseous fuels–oxygen mixtures

An experimental study on rotating detonation is presented in this paper. The study was focused on the possibility of using rotating detonation in a rocket engine. The research was divided into two parts: the first part was devoted to obtaining the initiation of rotating detonation in fuel–oxygen mixture; the second was aimed at determination of the range of propagation stability as a function of chamber pressure, composition, and geometry. Additionally, thrust and specific impulse were determined in the latter stage. In the paper, only rich mixture is described, because using such a composition in rocket combustion chambers maximizes the specific impulse and thrust. In the experiments, two kinds of geometry were examined: cylindrical and cylindrical-conic, the latter can be simulated by a simple aerospike nozzle. Methane, ethane, and propane were used as fuel. The pressure–time courses in the manifolds and in the chamber are presented. The thrust–time profile and detonation velocity calculated from measured pressure peaks are shown. To confirm the performance of a rocket engine with rotating detonation as a high energy gas generator, a model of a simple engine was designed, built, and tested. In the tests, the model of the engine was connected to the dump tank. This solution enables different environmental conditions from a range of flight from 16 km altitude to sea level to be simulated. The obtained specific impulse for pressure in the chamber of max. 1.2 bar and a small nozzle expansion ratio of about 3.5 was close to 1,500 m/s.

Heat Flux and Pressure Profiles in an Oxygen/Hydrogen Multielement Rocket Combustor

Proof of the predictive accuracy of models used in simulation of the high-pressure, high-heat flux environments inside operating rocket combustion chambers is limited by several factors including multiple phases, unsteadiness, high energy release, turbulence, and complex flow geometry. The severe environment also complicates the measurement of temperature, pressure, and chemical composition within the combustor. This paper describes an experiment to collect and analyze benchmark-quality validation data for a multi-element rocket combustor using liquid oxygen and gaseous hydrogen operating at chamber pressures from 300 to 1000 psia at mixture ratios from three to six. For these propellants, the stoichiometric mixture ratio is eight. The combustor was instrumented with 44 coaxial thermocouple pairs to measure axial and circumferential temperature fields and 8 pressure transducers to measure axial pressure profiles. The local heat flux at the inner wall of the combustion chamber computed at each temperature measurement location was found to vary between 4 and 20 Btu/in. 2 /s. The circumferential variation of heat flux seems to indicate mixing and radial outflow of combustion gases between injector elements, a phenomenon that is not observed in injector configurations with tighter spacing. Further, the axial pressure profiles and head-end temperature measurements suggest the presence of recirculation zones near the injector face.

연소시스템의 열음향 불안정 예측을 위한 Helmholtz Solver 개발

본 연구에서는 실제 로켓엔진 및 가스터빈용 연소기 내부의 열음향 불안정을 효과적으로 예측하기 위하여, 헬름홀츠 방정식과 시간지연모델을 이용한 3차원 유한요소법 해석코드를 개발하였다. 연소응답항에 의해 수치적으로 야기되는 비선형성은 반복법으로 선형화 하였으며, Arnoldi 방법을 사용하여 대용량 고유치 문제를 해석하였다. 해석결과인 복소각주파수와 음향 압력장을 통해 각 음향모드의 공진주파수, 진폭의 증폭/감쇠 여부 그리고 모드 형태를 예측할 수 있다. 이론해가 존재하는 두 가지 문제를 통해 출구 임피던스와 예혼합 화염이 종 방향 음향장에 미치는 영향에 대한 예측 정확도를 평가하였으며, 배플 유무에 따른 횡 방향 음향 모드의 주파수 변이를 상온 음향시험 결과와 비교/검증하였다. In order to effectively predict thermo-acoustic instability within real combustors of rocket engines and gas turbines, in the present study, the Helmholtz equation in conjunction with the time lag hypothesis is discretized by the finite element method on three-dimensional hybrid unstructured mesh. Numerical nonlinearity caused by the combustion response term is linearized by an iterative method, and the large-scale eigenvalue problem is solved by the Arnoldi method available in the ARPACK. As a consequence, the final solution of complex valued eigenfrequency and acoustic pressure field can be interpreted as resonant frequency, growth rate, and modal shape for acoustic modes of interest. The predictive capabilities of the present method have been validated against two academic problems with complex impedance boundary and premixed flame, as well as an ambient acoustic test for liquid rocket combustion chamber with/without baffle.

Femtotechnology: Design of the Strongest AB Matter for Aerospace

Aerospace, aviation particularly, need, in any era, the strongest and most thermostable materials available, often at nearly any price. The space elevator, space ships (especially during atmospheric reentry), rocket combustion chambers, thermally challenged engine surfaces, hypersonic aircraft materials are better now than any available, with undreamed performance as the reward if obtained. As shown in this research, the offered new material allows to greatly improve all characteristics of space ships, rockets, engines, and aircraft, and design new types space, propulsion, and aviation systems. At present the term “nanotechnology” is well known—in its ideal form, the flawless and completely controlled design of conventional molecular matter from molecules or atoms. Such power over nature would offer routine achievement of remarkable properties in conventional matter, and creation of metamaterials where the structure not the composition brings forth new powers of matter. But even this yet unachieved goal i...

Liquid rocket combustion chamber acoustic characterization

Over the last 40 years, many solid and liquid rocket motors have experienced combustion instabilities. Among other causes, there is the interaction of acoustic modes with the combustion and/or fluid dynamic processes inside the combustion chamber. Studies have been showing that, even if less than 1% of the available energy is diverted to an acoustic mode, combustion instability can be generated. On one hand, this instability can lead to ballistic pressure changes, couple with other propulsion systems such as guidance or thrust vector control, and in the worst case, cause motor structural failure. In this case, measures, applying acoustic techniques, must be taken to correct/minimize these influences on the combustion. The combustion chamber acoustic behavior in operating conditions can be estimated by considering its behavior in room conditions. In this way, acoustic tests can be easily performed, thus identifying the cavity modes. This paper describes the procedures to characterize the acoustic behavior in the inner cavity of four different configurations of a combustion chamber. Simple analytical models are used to calculate the acoustic resonance frequencies and these results are compared with acoustic natural frequencies measured at room conditions. Some comments about the measurement procedures are done, as well as the next steps for the continuity of this research. The analytical and experimental procedures results showed good agreement. However, limitations on high frequency band as well as in the identification of specific kinds of modes indicate that numerical methods able to model the real cavity geometry and an acoustic experimental modal analysis may be necessary for a more complete analysis. Future works shall also consider the presence of passive acoustic devices such as baffles and resonators capable of introducing damping and avoiding or limiting acoustic instabilities.

S0302-2-2 レーザー超音波可視化技術の開発とロケット燃焼器の非破壊検査への応用([S0302-2]超音波非破壊評価の高精度化(2))

S0302-2-2 レーザー超音波可視化技術の開発とロケット燃焼器の非破壊検査への応用([S0302-2]超音波非破壊評価の高精度化(2))

Degradation of space exposed surfaces by hypervelocity dust bombardment, and refractory materials for space

Abstract Dust particles with diameters below 100 μm represent an important part of the space environment. Objects like satellites or spacecrafts, are constantly bombarded with particles of cosmic velocities of 10 km/s and more. These hypervelocity impacts lead to evaporation of a large fraction of these particles and to the formation of craters on the material surfaces which exhibit diameters which are up to one order of magnitude larger than the impinging particles. This results in a remarkable materials degradation of space exposed surfaces which can considerably reduce the life time e.g. of semiconductor panels for the generation of energy, due to the cumulative effect of a multitude of such particles which hit a surface during a certain time. The particle population in space near to earth consists of two distinct groups: (a) Natural particles, micrometeorites: These particles are a natural part of the space environment and cannot be avoided. Most of them have so-called “chondritic” compositions and consist of Si, Al, Mg, Fe, Ca and O, an elemental makeup resembling the bulk composition of the original solar nebula from which our solar system and the planets formed 4.5 billion years ago. (b) Man made debris particles: These stem from various sources, e.g. exhausts from solid rocket fuels, paint chips, fragments from space crafts etc. These particles pose an increasing danger to space flight and their future production must be reduced by strict debris policies. The important question was now: What percentage of impact particles is man-made? Cosmic and man-made particles exhibit different chemical compositions. Hence, a micro-chemical analysis of particle impacts on satellite surfaces can be used to determine the particle ratio. For this purpose, the LDEF-experiment (Long Duration Exposure Facility) was launched by NASA in 1984 to study erosion phenomena on material surfaces (of more than 10,000 test samples) in a Low Earth Orbit (LEO)-environment. LDEF was retrieved only in 1990 because of the Challenger disaster which essentially delayed the retrieval up to the last moment before the experimental setup would have been destroyed upon entrance into the earth atmosphere. The demonstration of pronounced material degradation of space exposed materials logically leads to the question which materials are used in space technology. Plansee SE in Reutte/Tyrol has quite a tradition in the production of smaller rocket nozzles, e.g. satellite propulsion nozzles and general propulsion components made of Mo, W, Nb or Ta-based alloys. The W–Ag evaporation cooled alloy was among the earliest materials for rocket combustion chambers enduring ultra high temperatures due to partial evaporation of the silver which thereby cooled the chamber. Ni- and Fe-based ODS superalloys find application in durable TPS/hot structures of hypersonic vehicles, e.g. as honeycomb panels and fasteners of PM 1000/2000. These are PM-ferritic and nickel based oxide dispersion strengthened alloys. We have studied the significance of Yttrium depletion within the oxide scales of these alloys.

Acoustic characteristics of a rocket combustion chamber: Radial baffle effects

This paper describes methods used for determining the characteristic acoustic modes and frequencies of a liquid–propellant rocket–motor combustion chamber and effects of radial baffles on the chamber’s acoustic field. A multi-point sensing experimental setup, including stationary and moving sensors, was used to measure characteristic frequencies and mode shapes of a combustion chamber. A new technique based on the comparison of signal phase angles from stationary sensors to that of a moving sensor was used to map complex characteristic mode shapes of a combustor. A three-dimensional Helmholtz acoustic solver was also developed using an efficient finite volume approach for complex geometries to simulate the acoustic field inside a combustor. Using this approach the effects of the convergent section of the nozzle and the number of radial baffles on the chamber’s dominant acoustic modes with no mean flow were investigated. We have shown that the classical reduction of characteristic frequency of tangential modes caused by radial baffles is due to longitudinalization of tangential modes and is a function of the blade length and weakly dependent on the number of blades. Also, conjugate spinning modes are decoupled and do not spin in any baffled combustor, independent of the number of blades. On the other hand the converging nozzle section of a combustion chamber modifies pure longitudinal modes in the radial direction and pure tangential modes in the longitudinal direction. Existence of some mixed tangential–longitudinal modes in a combustor is dependent on the ratio of the nozzle throat diameter to the combustor head plate diameter.

Transient nonlinear optically-thick radiative–convective double-diffusive boundary layers in a Darcian porous medium adjacent to an impulsively started surface: Network simulation solutions

A boundary-layer model is described for the two-dimensional nonlinear transient thermal convection heat and mass transfer in an optically-thick fluid in a Darcian porous medium adjacent to an impulsively started vertical surface, in the presence of significant thermal radiation and buoyancy forces in an (X∗,Y∗,t∗) coordinate system. An algebraic approximation is employed to simplify the integro-differential equation of radiative transfer for unidirectional flux normal to the plate into the boundary-layer regime, by incorporating this flux term in the energy conservation equation. The conservation equations are non-dimensionalized into an (X,Y,T) coordinate system and solved using the Network Simulation Method (NSM), a robust numerical technique which demonstrates high efficiency and accuracy. The transient variation of non-dimensional streamwise velocity component (u) and temperature (T) and concentration (C) functions is computed for various selected values of Stark number (radiation–conduction interaction parameter) and Darcy number. Transient velocity (u) and steady-state local skin friction (τX) are also studied for various thermal Grashof number (Gr), species Grashof number (Gm), Schmidt number (Sc) and Stark number (N) values. These computations for the infinite permeability case (Da→∞) are compared with previous finite difference solutions [Prasad et al. Int J Therm Sci 2007;46(12):1251–8] and shown to be in excellent agreement. An increase in Darcy number is seen to accelerate the flow and boost velocity. A decrease in Stark number (corresponding to an increase in thermal radiation heat transfer contribution) is shown to increase the velocity values. Temperature function is observed to fall in value with a rise in Da and increase with decrease in N (corresponding to an increase in thermal radiation heat transfer contribution). Applications of the study include rocket combustion chambers, astrophysical flows, spacecraft thermal fluid dynamics in debris-laden environments (cosmic dust), heat transfer in forest fire spread, geochemical contamination and ceramic materials processing.

Film Cooling of Accelerated Flow in a Subscale Combustion Chamber

Experimental investigations have been carried out to examine film cooling effectiveness of an accelerated hot gas in a subscale rocket combustion chamber. In support of future first-stage high-performance rocket combustion chambers, a Vulcain2-like test case has been examined with combustion pressure levels up to 12 MPa. The effectiveness of an almost tagentially injected film of hydrogen with an initial temperature of approximately 280 K has been determined. Axial distributions of temperature were measured inside the copper liner as well as on the chamber surface in the convergent and divergent parts of the nozzle segment. An existing film cooling model has been modified for application in a combined convective and filmcooled combustion chamber with an accelerated hot gas. The new model predicts film cooling effectiveness at different combustion-chamber pressures and film blowing rates at sub-, trans-, and supersonic conditions.

Flame Stabilization in High-Pressure Liquid Oxygen/Methane Rocket Engine Combustion

Flame stabilization in the near-injector region of a subscale liquid propellant rocket combustion chamber has been investigated using optical diagnostics. Several different single shear coaxial injectors have been used to inject liquid oxygen (LOX) and methane at relevant operating conditions. Three main operating points cover the range of sub-, near-, and superciritical conditions with respect to the thermodynamic cirtical point of oxygen. Whereas liquid temperatures of about 270 K. Similar to previous investigations performed with LOX/H2 combustion, spontaneous OH chemiluminescene has been detected to characterize the flame anchoring zone near the LOX post tip. Theoretical considerations have indicated that the binary mixture of oxygen and methane shows and entirely different behaviour compared with the oxygen/hydrogen system. This is believed to have an influence on the spray evolution and mixing characteristics at supercritical conditions. The experimental investigation includes both the ignition transient a well as the steady-state operating points. It has been detected that the LOX/CH4 flame shows very similar characteristics in comparison with LOX/H2 flame at similar operating conditions. Critical flame stabilization at the startup transient has been found during all hot runs. At steady-state conditions, the influence of the injection parameters on the flame shape is comparable to previous LOX/Hydrogen investigations.

Experimental Investigation of Film Cooling with Tangential Slot Injection in a LOX/CH4 Subscale Rocket Combustion Chamber

Within the frame of a broader activity towards research into the application of methane in cryogenic liquid rocket engines, the efficiency of film cooling was studied in a LOX/CH4 fired subscale sized model combustor. Aiming at booster as well as upper stage applications, the combustion pressure levels have been varied between 4 MPa and 7 MPa. The effectiveness of the ambient temperature film was determined in axial and circumferential directions by measuring temperature gradients in the copper liner material. The experiments revealed remarkable circumferential differences of the film cooling efficiency which remain existent far downstream. However, circumferential film cooling varieties are more pronounced at close proximity of the point of film coolant injection.

Heat flux measurement in a high enthalpy plasma flow

It is a widely used approach to measure heat flux in harsh environments like high enthalpy plasma flows, fusion plasma and rocket motor combustion chambers based on solving the inverse heat conduction problem in a semi-infinite environment. This approach strongly depends on model parameters and geometrical aspects of the sensor design. In this work the surface heat flux is determined by solving the inverse heat conduction problem using an identified system as a direct model. The identification of the system is performed using calibration measurements with modern laser technique and advanced data handling. The results of the identified thermo-physical system show that a non-integer model appears most adapted to this particular problem. It is concluded that the new method improves the heat flux sensor significantly and furthermore extend its application to very short measurement times.

Thermal sensor calibration using non-integer system identification

A widely used approach to measuring heat flux in harsh environments such as high enthalpy plasma flows, fusion plasma and rocket motor combustion chambers is based on solving the inverse heat conduction problem in a semi-infinite environment. This approach strongly depends on model parameters and geometrical aspects of the sensor design. In the present paper the surface heat flux is determined by solving the inverse heat conduction problem using an a priori identified system as a direct model within the estimation process. Calibration measurements using modern laser technique are used to identify the physical thermal system, where it turns out that a non-integer model appears most suitable to this particular problem. It is concluded that with the new method the heat flux sensor is significantly improved. Furthermore, its applicability is extended to very short measurement times.

Nonlinear rocket motor stability prediction: Limit amplitude, triggering, and mean pressure shift

High-amplitude pressure oscillations in solid propellant rocket motor combustion chambers display nonlinear effects including: (1) limit cycle behavior in which the fluctuations may dwell for a considerable period of time near their peak amplitude, (2) elevated mean chamber pressure (DC shift), and (3) a triggering amplitude above which pulsing will cause an apparently stable system to transition to violent oscillations. Along with the obvious undesirable vibrations, these features constitute the most damaging impact of combustion instability on system reliability and structural integrity. The physical mechanisms behind these phenomena and their relationship to motor geometry and physical parameters must, therefore, be fully understood if instability is to be avoided in the design process, or if effective corrective measures must be devised during system development. Predictive algorithms now in use have limited ability to characterize the actual time evolution of the oscillations, and they do not supply the motor designer with information regarding peak amplitudes or the associated critical triggering amplitudes. A pivotal missing element is the ability to predict the mean pressure shift; clearly, the designer requires information regarding the maximum chamber pressure that might be experienced during motor operation. In this paper, a comprehensive nonlinear combustion instability model is described that supplies vital information. The central role played by steep-fronted waves is emphasized. The resulting algorithm provides both detailed physical models of nonlinear instability phenomena and the critically needed predictive capability. In particular, the origin of the DC shift is revealed.

Vaporization Lengths of Hydrazine Fuels Burning with NTO

The performances of single hydrazines and hydrazine mixtures burning with NTO in a bipropellant rocket combustion chamber are compared. Systems using single hydrazines (N2H4, MMH, UDMH) and hydrazine mixtures (MMH=N2H4 and UDMH=N2H4) were simulated with total mass flow rates of 45:29 g=s and 89:25 g=s, respectively, in a chamber with diameter 40 mm. A one-dimensional mathematical model was developed, assuming thatcombustionandturbulentmixtureprocesses arecontrolledbydropletvaporization,andthatdropletdiameters follow the Rosin–Rammler distribution function with a finite number of diameters. The effects of drag, thermal expansion, transient vaporization, and heat transfer by convection and radiation to the droplets and to the chamber wallswereconsidered.Theinfluencesofchamberpressure,equivalenceratio,initialgastemperature,numberofsize classes, Sauter mean diameter, and parameters of the Rosin–Rammler distributions on vaporization lengths of dropletsareanalyzed.Profilesofnormalizeddiameters,temperatures,velocities,heatlosses,vaporizationrates,and other flow properties along the chamber are described. Nomenclature A = area or constant of Antoine’s equation B = constant of Antoine’s equation BM = transfer number C = constant of Antoine’s equation CD = drag coefficient Cp = specific heat at constant pressure D = diameter D32 = Sauter mean diameter F = drag f = mixture ratio h = specific enthalpy hfg = vaporization enthalpy ~ h � = convection coefficient for droplet heating

Rocket Propellant Characteristics of Silanes/O2

We present a theoretical analysis of the applicability of various silicon hydrides (silanes) as rocket propellants. Thermodynamic data sets for 12 different silanes ranging from monosilane SiH4 to pentasilane Si5H12 were generated based on existing data. They were derived from empirically corrected thermochemical ab initio calculations and used in conjunction with the well-known computer packages NASA CEA2 and CHEMKIN to calculate equilibrium compositions and rocket performance indices. The thermodynamic datasets are now publicly available online for future calculations. Chemical equilibrium combustion compositions and adiabatic flame temperatures in rocket combustion chambers and the specific impulses at the nozzle exit assuming shifting equilibrium were calculated for several silane fuels using oxygen as oxidizer. The computational results created by CEA2 and CHEMKIN are in very good agreement to each other and are compared to conventional fuels such as liquid hydrogen (LH), hydrocarbons and hydrazines. Compared to alkane fuels, the combustion of silanes with oxygen-despite the much higher average molecular mass of the combustion products-results in similar maximum specific impulses and adiabatic flame temperatures due to the high positive values of the heats of formation of the silanes, indicating that silanes may be interesting candidates as rocket propulsion fuels.

Numerical study of liquid film cooling in a rocket combustion chamber

A numerical study is reported to investigate the liquid film cooling in a rocket combustion chamber. Mass, momentum and heat transfer characteristics through the interface are considered in detail. A marching procedure is employed for solution of the respective governing equations for the liquid film and gas stream together. The standard turbulence k– ε model is used to simulate the turbulence gas flow and a modified van Driest model is adopted to simulate the turbulent liquid film flow. Radiation of gas stream is also considered and simulated with the flux model. Downstream of the liquid film the gaseous film cooling is numerically studied simultaneously. Results are presented for a mixed gases–water system under different condition. Various effects on the liquid film length are examined in detail. There is a good agreement between the numerical prediction and experimental result on the liquid film length.