Showing papers in "Journal of Propulsion and Power in 2002"

Thermodynamic Cycle Analysis of Pulse Detonation Engines

Abstract: Pulse detonation engines (PDEs) are currently attracting considerable research and development attention because they promise performance improvements over existing airbreathing propulsion devices. Because of their inherently unsteady behavior, it has been difficult to conveniently classify and evaluate them relative to their steady-state counterparts. Consequently, most PDE studies employ unsteady gasdynamic calculations to determine the instantaneous pressures and forces acting on the surfaces of the device and integrate them over a cycle to determine thrust performance. A classical, closed thermodynamic cycle analysis of the PDE that is independent of time is presented. The most important result is the thermal efficiency of the PDE cycle, or the fraction of the heating value of the fuel that is converted to work that can be used to produce thrust. The cycle thermal efficiency is then used to find all of the traditional propulsion performance measures. The benefits of this approach are 1) that the fundamental processes incorporated in PDEs are clarified; 2) that direct, quantitative comparisons with other cycles (e.g., Brayton or Humphrey) are easily made; 3) that the influence of the entire ranges of the main parameters that influence PDE performance are easily explored; 4) that the ideal or upper limit of PDE performance capability is quantitatively established; and 5) that this analysis provides a basic building block for more complex PDE cycles. A comparison of cycle performance is made for ideal and real PDE, Brayton, and Humphrey cycles, utilizing realistic component loss models. The results show that the real PDE cycle has better performance than the real Brayton cycle only for flight Mach numbers less than about 3, or cycle static temperature ratios less than about 3. For flight Mach numbers greater than 3, the real Brayton cycle has better performance, and the real Humphrey cycle is an overoptimistic (and unnecessary) surrogate for the real PDE cycle.

Combustion of Liquefying Hybrid Propellants: Part 1, General Theory

Abstract: In this paper classical hybrid combustion theory is generalized to solid fuels that form a liquid layer on their burning surface. For several classes of liquefying fuels, the layer is hydrodynamically unstable in a gas e ow environment leading to substantial droplet entrainment into the gas stream. The susceptibility of a given fuel to this shear driven instability increases with decreasing viscosity and surface tension of the melt layer. The entrainment mass transfer, which acts in addition to the conventional gasie cation mechanism, is not affected by the blocking phenomenon induced by blowing from the surface. For practical oxidizer e ux levels encountered in hybrid rocket applications, droplet entrainment can dominate direct gasie cation. Such liquefying fuels can exhibit greatly increased surface regression rates compared to classical materials such as Hydroxyl Terminated Polybutadiene. One application of the theory is to solid cryogenic hybrids, which utilize frozen materials for the solid propellant. The theory successfully predicts why high regression rates are observed in tests of cryogenic solid pentane, solid methane, and solid oxygen. In addition, the theory explains the dependence of the burning rates of other tested cryogenic materials on the physical properties of the liquid layer. The theory also leads to the conclusion that certain noncryogenic materials such as parafe n and Polyethylene waxes will also exhibit high regression rates. This important result is cone rmed by lab scale tests performed at Stanford University on a high melting point parafe n wax.

Experimental Investigation of Limit Cycle Oscillations in an Unstable Gas Turbine Combustor

Abstract: This paper describes an experimental investigation of the characteristics of limit cycle oscillations in an unstable gas turbine combustor simulator. This investigation was performed to improve the current understanding of the nonlinear processes controlling these oscillations. Such an understanding is needed in order to predict instability amplitudes, to aid in correlating data, and to develop and optimize active control methodologies. The paper first describes an analysis of the statistical and temporal features of the limit cycle pressure oscillations, and discusses the role of system nonlinearities upon these oscillations. Next, it discusses the important role that the combustor inlet velocity plays in determining the amplitude of the limit cycle oscillations. The paper also presents data illustrating the characteristics of the combustor's transition from stable to unstable operation, and shows that these characteristics can be used to predict the occurrence of nonlinear phenomenon (e.g., hysteresis) that are often observed in unstable combustors. Finally, it is shown that inherent noise in the system can strongly affect the limit cycles, and may even be responsible for causing the combustor to become unstable under nominally stable conditions. The paper concludes with a discussion of the implications of these results on the current understanding of self-excited, combustion driven oscillations in lean, premixed gas turbine combustors.

Study of the High-Temperature Autoignition of n-Alkane/O/Ar Mixtures

Abstract: Ignition time measurements of propane, n-butane, n-heptane, and n-decane have been studied behind reflected shock waves over the temperature range of 1300-1700 K and pressure range of 1-6 atm. The test mixture compositionvaried from approximately 2-20% O 2 , and the equivalence ratio ranged from 0.5 to 2.0. To determine more precisely the fuel mole fraction of the test mixture, a new technique has been employed in which a 3.39-μm HeNe laser and multiple-pass setup is utilized to measure the fuel in situ by absorption. Ignition delay times were measured at the shock tube endwall by a CH emission diagnostic (431 nm) that viewed the shock-heated mixture through a window in the endwall. This enabled the ignition time at the unperturbed endwall conditions to be determined accurately, thereby avoiding problems inherent in measuring ignition times from the shock tube sidewall. A parametric study of the experimental data reveals marked similarity of the ignition delay time characteristics among these four n-alkanes, and a unique correlation is presented in which the stoichiometric ignition time data for all four n-alkanes has been correlated into a single expression with an R 2 value of 0.992: Τ=9.4×10 - 1 2 P - 0 . 5 5 X O 2 - 0 . 6 3 C - 0 . 5 0 exp(46,550/RT) where the ignition time is in seconds, pressure in atmospheres, the activation energy in calories per mole, X O 2 is the mole fraction of oxygen in the test mixture, and C is the number of carbons atoms in the n-alkane. Comparisons to past ignition time studies and detailed kinetic mechanisms further validate the correlations presented here.

Direct Experimental Impulse Measurements for Detonations and Deflagrations

Abstract: Direct impulse measurements were carried out by using a ballistic pendulum arrangement for detonations and deflagrations in a tube closed at one end. Three tubes of different lengths and inner diameters were tested with stoichiometric propane– and ethylene–oxygen–nitrogen mixtures. Results were obtained as a function of initial pressure and percent diluent. The experimental results were compared to predictions from an analytical model and generally agreed to within 15% (Wintenberger, E., Austin, J., Cooper, M., Jackson, S., and Shepherd, J. E., “Analytical Model for the Impulse of a Single-Cycle Pulse Detonation Engine, AIAA Paper 2001–3811, July 2001). The effect of internal obstacles on the transition from deflagration to detonation was studied. Three different extensions were tested to investigate the effect of exit conditions on the ballistic impulse for stoichiometric ethylene–oxygen–nitrogen mixtures as a function of initial pressure and percent diluent.

Combustion of Liquefying Hybrid Propellants: Part 2, Stability of Liquid Films

Abstract: The stability of a liquid layer under strong blowing and subjected to large shear forces is investigated. This case is of practical importance for application to theregression rate estimation of liquefying hybrid rocket fuels such as solid cryogenichybrids. An Orr ‐Sommerfeld equation forthelinearstability of the liquid ‐gas interface is derived, and an exact solution is found for a linear base velocity proe le. The exact solution for the liquid phase is coupled with thelinearized gas-phase response with appropriate boundary conditions at the interfaceto give an eigenvalue problem for the linear stability of the layer. The results for liquid layer Reynolds numbers of practical interest (50‐300) show the existence of a range of unstable wave numbers. It is observed that both the most amplie ed wave number and the maximum amplie cation increases with the liquid Reynolds number. It is also discovered that increasing surface tension and liquid viscosity have a stabilizing effect on the e lm. This prediction is supported by experimental results showing fast burning rates for low-viscosity fuels such as solid cryogenic pentane and noncryogenicwax. Finally, thestability theory is applied to the classical polymeric hybrid propellants that burn by forming a melt layer. Because the melt layers of these polymeric materials are highly viscous, they can not sustain thin e lm instabilities.

Control of Low-Pressure Turbine Separation Using Vortex-Generator Jets

Abstract: The application of vortex-generator jets to control separation on the suction surface of a low-pressure turbine blade is reported. Blade Reynolds numbers in the experimental, linear turbine cascade match those for high-altitude operation of many aircraft gas-turbine engines, as well as the last stages of industrial ground-based gas turbines. Results are presented for steady blowing at jet blowing ratios from zero to four and at several chordwise positions and two freestream turbulence levels. Findings show that above a minimum blowing ratio, which is dependant on the injection location, the pressure loss in the modified blade's wake is reduced by a factor of between two and three. Boundary-layer traverses show that separation is almost completely eliminated with the application of blowing. No significant deleterious effects of vortex-generator jets are observed at higher (nonseparating) Reynolds numbers. The addition of 4% freestream turbulence to the cascade freestream lowers the separation Reynolds number of the turbine blade studied, but does not eliminate the effectiveness of the control technique. The vortex-generator jet control strategy is demonstrated to be a viable technique for low-pressure turbine separation control.

Solid-Fuel Regression Rate Behavior of Vortex Hybrid Rocket Engines

Abstract: Aseriesofstaticenginee ringswereconductedtoinvestigatethesolid-fuelregressionratebehaviorandoperating characteristicsofvortexhybridrocketengines.Thevortexhybridenginecone gurationischaracterizedbyacoaxial, coswirling, countere owing vortex combustion e eld in a cylindrical fuel port. To generate this e owe eld, oxidizer is injected through a swirl injector located between the aft end of the fuel grain and the inlet to the converging portion of the exit nozzle. Test e rings with thrusts up to 960 N were conducted with gaseous oxygen and hydroxylterminated polybutadiene solid fuel. Average fuel regression rates up to seven times larger than those in similar classical hybrids were measured. Empirical correlations were developed to describe accurately the experimental regression rates over more than an order of magnitude variation in mass e ux. In addition to local mass e ux and oxygeninjectionvelocity,geometricenginevariables,suchasenginecontractionratioand length-to-diameterratio, had a signie cant ine uence on the measured regression rates. Nondimensional regression rate and heat transfer correlations were also developed. Throttling and restart capability were demonstrated. Nomenclature A = cross-sectional area, cm 2 B = blowing parameter cp = isobaric specie c heat, J/kg ¢K D = port diameter, cm Ea = activation energy, kcal/mole G = local mass e ux, kg/m 2 ¢s Ginj = injection mass velocity, kg/m 2 ¢s GO = gaseous oxygen mass e ux, kg/m 2 ¢s

Analysis and Application of Chorochronic Periodicity in Turbomachinery Rotor/Stator Interaction Computations

Abstract: A methodology for predicting the three-dimensional unsteady aerodynamics of the interaction between two turbomachinery blade rows that are in relative angular motion with one another is described. In this case, the kinematics of the blades introduce a chorochronic (space‐time) periodicity. This periodicity is analyzed in detail, and a mathematically straightforward methodology, based on Fourier series in µ (azimuth) and t (time), is presented for treating the interface between the two rows. These results are implemented in a computational method solving the three-dimensional Favre ‐Reynolds-averaged Navier ‐Stokes equations, with a near-wall wall-normalfree Reynolds-stress model. Only one blade passage per blade row is discretized. At the pitchwise boundaries, phase-lagged periodicity is applied using Fourier series in time. Both the pitchwise boundaries time harmonics and the interface chorochronic tµ harmonics are updated using a low-storage moving-averages technique. Computational results are presented and compared with measurements for a 1 1 -stage turbine, where the two stators have the same number of blades enabling the use of chorochronic periodicity. Sample results are also presented for a transonic inlet guide vane/rotor interaction, illustrating the ability of the interface treatment to handle shock waves.

Analysis of Dual-Mode Hydrocarbon Scramjet Operation at Mach 4-6.5

Abstract: Reynolds averaged Navier ‐Stokes calculations have been performed for a U.S. Air Force Research Laboratory/ Aerospace Propulsion Ofe ce scramjet combustor designed for Mach 4.0 ‐6.5 e ight. The combustor e owpath is unique in that it is entirely free of e ow obstructions with fuel injection from wall-mounted injection ports and e ameholding established by means of a recessed cavity. Calculations were performed at the minimum (Mach 4.0) andmaximum(Mach6.5)e ightdesign conditions.Thecombustoroperatedin dualmodeattheMach4.0condition. The precombustion shock train formed a region of low-momentum/separated e ow adjacent to the combustor side wall. This proved to be a primary source of e ameholding, with the recessed cavity adding additional e ameholding support. The e ow was notthermally choked at theMach 6.5 condition,resulting in very little upstream interaction. The mixing process at the Mach 4.0 e ight condition was considerably more efe cient than that seen at the Mach 6.5 condition, due primarily to the shock-induced e ow distortion and larger residence time. Even with the reduced mixing levels predicted at the Mach 6.5 condition, the combustion efe ciency was comparable to that achieved at the Mach 4.0 condition. The solutions obtained for dual-mode operation were particularly sensitive to choice of turbulence model and values specie ed for the turbulent Prandtl and Schmidt numbers. Overall, the solution sensitivity to grid resolution was small relative to the solution sensitivity to modeling uncertainties.

Appearance of Restricted Shock Separation in Rocket Nozzles

Abstract: Uncontrolled flow separation in nozzles of rocket engines is not desired because it can lead to dangerous lateral forces. Different origins for side loads were identified in the past. Meanwhile, it is proven that in thrust-optimized or parabolic nozzles, a major side load occurs as a result of the transition of separation pattern from free shock separation to restricted shock separation and vice versa. Reasons for the transition between the separation patterns are discussed, and the cap-shock pattern, which is identified to be the cause of this transition, is closely analyzed. It turns out that this pattern can be interpreted as an inverse Mach reflection of the internal shock at the nozzle axis. To prove the transition effect as main side-load driver, a subscale test campaign has been performed. Two different nozzle contours, a thrust-optimized and a truncated ideal nozzle with equal performance data, were tested. Highest side loads were measured in the thrust-optimized nozzle, when the separation pattern changes from free to restricted shock separation. Side loads measured in the truncated ideal nozzle were only about one-third as high as in the thrust-optimized nozzle.

Advanced Performance Modeling of Experimental Laser Lightcraft

Abstract: A computational plasma aerodynamics model is developed to study the performance of a laser-propelled lightcraft. The computational methodology is based on a time-accurate, multi-dimensional, finite volume, chemically reacting, unstructured grid pressure-based formulation. The underlying physics are modeled using a building-block approach. The physics modeled include nonequilibrium thermodynamics, nonequilibrium air-plasma finite rate kinetics, specular ray tracing, laser beam energy absorption and refraction by plasma, nonequilibrium plasma radiation, and plasma resonance. A series of transient computations are performed at several laser pulse energy levels and the simulated physics are discussed and compared with those of tests and literatures. The computed impulses and coupling coefficients for the lightcraft compared reasonably well with those of tests conducted on a pendulum apparatus.

Burning Aluminum Particles Inside a Laboratory-Scale Solid Rocket Motor

Abstract: Single and agglomerated aluminum droplets were studied in a solid rocket motor test chamber with optical access to the internal flow at 6-22 atm and 2300 K. The chamber was pressurized by burning a main grain ammonium perchlorate/hydroxyl-terminated poly-butadiene propellant, and the burning aluminum droplets were generated by a smaller aluminized solid propellant sample, center mounted in the flow. A 35-mm camera was used with a chopper wheel to give droplet flame diameter vs time measurements of the burning droplets in flight, from which burning rate laws were developed. A high-speed video charge-coupled device with high-magnification optics imaged the flame/smoke cloud surrounding the burning liquid droplets. The intensity profiles of the droplet images were deconvoluted using an Abel inversion to give true intensity profiles. Both single and agglomerated droplets were studied, where agglomerates comprise hundreds of parent particles or more. The Abel inversions show that the relative smoke cloud size is not constant with diameter, but instead grows as the droplet shrinks, by ∼D - 0 . 5 , for both the single and agglomerated droplets. Measured diameter trajectories show that, for single droplets, the mean diameter law is D 0 . 7 5 = D 0 . 7 5 0 - 8.t, and, for agglomerated droplets, D 1 . 0 = D 1 . 0 0 - 20.t. For both single and agglomerated droplets, the burning rate slope k did not change significantly for the chamber pressure range studied.

Development of Advanced Industrial Furnace Using Highly Preheated Combustion Air

Abstract: Results are presented on the thermal and chemical characteristics of flames using high-temperature combustion air and liquified petroleum gas (LPG) as the fuel. The stability limits of these flames are extremely wide as compared to any other method of flame stabilization. This study is part of the Japan national project directed to develop advanced industrial furnace designs that provide approximately 30% energy savings and hence CO 2 reduction, 30% reduction in the furnace size, and 25% reduction of pollutants including NO x as compared to current designs. The objective here is to establish conditions that permit significant reduction in energy consumption, high efficiency, and low pollution from a range of furnaces. Data have been obtained on mean flame temperature and temperature fluctuations, flame emission spectra, emission intensity of C 2 and CH species from within the flames, and overall pollutant emission from the flames. The uniformity of temperature in the furnace was found to be far greater with low oxygen concentration combustion air preheated to 1000°C as compared to that obtained with roomtemperature air or that found in conventional flames. Emission of NO x and CO was much lower with combustion air preheated to high temperatures with low oxygen concentration. The chemiluminescence intensity of CH and C 2 radicals is significantly affected by the preheat temperature of the combustion air and oxygen concentration in the oxidant. The flame signatures revealed important flame characteristics under high-temperature air combustion conditions. The advantages of utilizing highly preheated combustion air (in excess of 1000°C) in various types of furnaces are given. The new and advanced furnace design utilizes high-efficiency regenerators and behaves essentially as a well-stirred reactor with uniform thermal and chemical characteristics. Because each furnace design requires unique features, it is imperative that each furnace must be optimized to satisfy the functional requirements of the furnace. In this paper a relatively simple diagnostic methodology is presented, which assists in a rational furnace design and optimization process.

Hydrogen Peroxide Decomposition on Various Supported Catalysts Effect of Stabilizers

Abstract: The decomposition of hydrogen peroxide (H 2 O 2 ) has been studied on various catalysts (platinum supported on silica; silver, iridium, platinum-tin or manganese oxides supported on alumina). The experiments were performed using two reactors: 1) a conventional constant pressure reactor for the determination of the volume increase vs time using diluted H 2 O 2 solutions; 2) a constant volume reactor to measure the pressure increase using more concentrated solutions. The first reactor leads to the determination of the kinetic order of the reaction, to the comparison of the activities of the different samples, and to the characterization of the influence of some stabilizers of H 2 O 2 solutions on the catalytic activity. Two kinetic orders were found, depending on the catalyst: a zero order and a first order. The shape of the catalysts samples is an important parameter, with powders always being more reactive than grains and pellets. The catalyst activities are sorted as follows: Pt-Sn/Al 2 O 3 < Ir/Al 2 O 3 < Pt/SiO 2 < MnO x /Al 2 O 3 < Ag/Al 2 O 3 . The presence of pyrophosphate stabilizer leads to a loss of activity mainly as a result of passivation in the case of MnO x -supported samples, whereas the presence of stannate increases slightly the activity of silver and displays no influence on manganese samples.

Gas Turbine Engine and Sensor Fault Diagnosis Using Optimization Techniques

Abstract: A diagnostic system is described for performance analysis of gas turbine engine components and sensors. The system estimates the performance parameters expressing the fault condition of the engine components in the presence of measurement noise and biases. The measurement uncertainty is supposed to affect even the parameters setting the operating condition of the engine. Estimation is performed through bptimization of an objective function by means of an ad hoc genetic algorithm. The genetic algorithm uses an accurate nonlinear steady-state performance model of the engine. The only statistical assumption required by the technique concerns the measurement noise and the maximum allowed number of faulty sensors and engine components, which is enclosed as a constraint. The technique has been thoroughly tested with the model of a low bypass ratio turbofan, and the results show the high level of accuracy achieved.

History of U.S. Navy Ramjet, Scramjet, and Mixed-Cycle Propulsion Development

Abstract: A history of high-speed airbreathing propulsion ramjet engines and their respective vehicle and weapon systems developed under the support of the U.S. Navy is presented. These include surface- and air-launched subsonic combustion ramjets, supersonic combustion ramjets (scramjets), and mixed-cycle ramjet/scramjet/rocket engines intended primarily for missile applications for flight speeds from Mach 2 to Mach 8. A summary of the development of the joint Department of Defense/NASA-sponsored National Aerospace Plane is also presented.

Shape optimization of supersonic turbines using global approximation methods

Abstract: There is growing interest to adopt supersonic turbines for rocket propulsion. However, this technology has not been actively investigated in the United States for the last three decades. To aid design improvement, a global optimization framework combining the radial-basis neural network (NN) and the polynomial response surface (RS) method is constructed for shape optimization of a two-stage supersonic turbine, involving O(10) design variables. The design of the experiment approach is adopted to reduce the data size needed by the optimization task. The combined NN and RS techniques are employed. A major merit of the RS approach is that it enables one to revise the design space to perform multiple optimization cycles. This benefit is realized when an optimal design approaches the boundary of a predefined design space. Furthermore, by inspecting the influence of each design variable, one can also gain insight into the existence of multiple design choices and select the optimum design based on other factors such as stress and materials consideration.

Combustion Instability Active Control Using Periodic Fuel Injection

Abstract: Active control using periodic fuel injection has the potential of suppressing combustion instability without radically changing the engine design or sacrificing performance A study is carried out of optimal model-based control of combustion instability using fuel injection The model developed is physically based and includes the acoustics, the heat-release dynamics, their coupling, and the injection dynamics A heat-release model with fluctuations in the flame surface area, as well as in the equivalence ratio, is derived It is shown that area fluctuations coupled with the velocity fluctuations drive longitudinal modes to resonance caused by phase-lag dynamics, whereas equivalence ratio fluctuations can destabilize both longitudinal and bulk modes caused by time-delay dynamics Comparisons are made between the model predictions and several experimental rigs The dynamics of proportional and two-position (on-off) fuel injectors are included in the model When the overall model is used, two different control designs are proposed The first is an linear quadratic Gaussian/loop transfer recovery controller, where the time-delay effect is ignored, and the second is a positive forecast controller, which explicitly accounts for the delay Injection at 1) the burning zone and 2) farther upstream is considered The characteristics of fuel injectors including bandwidth, authority (pulsed-fuel flow rate), and whether it applies a proportional or a two-position (on-off) injection are discussed We show that increasing authority and increasing bandwidth result in improved performance Injection at location 2 compared to location 1 results in a tradeoff between improved mixing and increased time delay It is also noted that proportional injection is more successful than on-off injection because the former can modulate both amplitude and phase of the control fuel

Magnetohydrodynamic Power Extraction from Cold Hypersonic Airflows with External Ionizers

Abstract: A novel concept of hypersonic cold-air magnetohydrodynamics (MHD) power generators with ionization by electron beams is analyzed. Electron beams are shown to allow control and stable operation of MHD channels in cold high-speed flows. To avoid excessive energy cost of ionization and damage to beam-injection foils, electron beam current densities should be restricted to a few milliamperes per square centimeter. This reduces the conductivity in electron beam sustained MHD channels compared with that in conventional MHD generators, restricting performance and calling for very strong magnetic fields and high Hall parameters. The high Hall parameters cause ion slip and near-anode phenomena to become first-order issues. Example one-dimensional calculations of hypersonic power generator performance appear to be promising. Possible problems that could be caused by hypersonic boundary layers and electrode sheaths, including anode sheath instability and ways to avoid it, are also discussed.

Improved Mean-flow Solution for Slab Rocket Motors with Regressing Walls

Abstract: The Navier‐Stokes equations are solved to obtain an approximate description of the mean e ow in a slab rocket motor with two evenly regressing walls. The scope is limited to two-dimensional incompressible and chemically nonreactive viscous e ow. The transformed governing equation is solved numerically, using e nite differences, and asymptotically, using variation of parameters and small parameter perturbations in the blowing Reynolds number R. Results are correlated and compared via variations in R and the dimensionless wall regression rate. For hard blowing and moderate regression rates the effect of wall motion on the velocity is found to be small. Conversely,forfast-burningpropellants,suchasthosebeing developed forhigh-accelerationvehicles,regressioneffects seem ine uential. Inclusion of viscous dissipation is also found to be important in assessing the total e ow vorticity, especially when R<102. The current geometric cone guration is relevant to motor simulations using ducted channels with porous walls. For validation purposes comparisons with numerical solutions are carried out alongside end-process verie cations. Because the resulting model incorporates viscosity and wall motion, it allows for an improved description of the unsteady acoustico-vortical solution whose assessment is strongly ine uenced by the mean e ow.

Ethylene Ignition and Detonation Chemistry, Part 1: Detailed Modeling and Experimental Comparison

Abstract: A numerical investigation is reported on the chemical kinetics of C 2 H 4 ignition and detonation in oxygen-diluent mixtures. Conditions addressed cover initial (postshock) temperatures between 1000 and 2500 K, pressures between 0.5 and 100 bar and equivalence ratios between 0.5 and 2. Attention is paid to histories of species concentrations and temperature, to autoignition times, and to subsequent heat release that is relevant to detonation structure. A detailed mechanism is proposed consisting of 148 reversible elementary reactions among 34 chemical species. This mechanism is shown to provide good agreement with most ignition times measured in shock tubes and with burning-velocity measurements available in the literature. The results provide a basis for developing simplified descriptions that can be used in multidimensional detonation studies for applications to propulsion devices.

Model of aluminum agglomerate evolution in combustion products of solid rocket propellant

Abstract: A mathematical model was developed for agglomerate evolution in the combustion products of solid rocket propellant. The model is based on the principle of dividing the complex process into relatively simple phenomena with sequential synthesis of their descriptions. The model includes description of many processes observed experimentally. They are 1 ) metal combustion, 2 ) chemical interaction between metal and oxide in agglomerates, 3) change in structure of agglomerates, and 4 ) interaction of agglomerates and carrier e ow. Quite high quality of the model was proved by satisfactory agreement between calculation results and experimental data.

Effect of nozzle cavity on resonance in large SRM : theoretical Modeling

Abstract: Cold gas experiments are used to study the vortex-nozzle interaction, which drives thrust pulsation in solid-rocket motors. The experiments carried out in an axial flow model clearly demonstrate coupling of vortex shedding with acoustical longitudinal resonances of the combustion chamber as observed in actual motors. The amplitudes of the pressure fluctuations correspond to one-thousandth of the static pressure, which is the order of magnitude of the observed pulsations in rocket motors. Experiments show that the cavity formed around the nozzle inlet during combustion is crucial. The pulsation level is proportional to the volume of the cavity. Theory predicts this relationship if we assume vortex-nozzle interaction to be the main source of sound. The proposed analytical model does, however, overestimate the pulsation level by an order of magnitude.

The Constant Volume Limit of Pulsed Propulsion for a Constant Gamma Ideal Gas

Abstract: The constant volume (CV) limit of pulsed propulsion was explored theoretically, where the combustion chamber was approximated as being time-varying but spatially uniform, while the nozzle flow was approximated as being one dimensional but quasi-steady. The CV limit was explored for the isentropic blow down of a constant y ideal gas for fixed expansion ratios and for variable expansion ratios which were adjusted to match the pressure ratio at all times. The CV limit calculations were notable in that all the fixed expansion ratio results could be expressed as analytical solutions. The CV limit was then compared with two relevant constant pressure (CP) cycles. Among the several conclusions were that using a fixed expansion ratio nozzle during a CV blow down did not exert more than a 3% performance penalty over using a variable expansion ratio nozzle as long as the fixed expansion ratio nozzle was optimized for the blow down. Another major conclusion was that, while the impulse produced by a CV device could significantly exceed that of a CP device operating at the fill pressure of the CV device at elevated ambient pressures (e.g. , I atm), the impulse produced by a CP device could actually exceed, albeit only slightly, the impulse produced by a CV device when the ambient pressure was near zero, such as would occur in the near vacuum conditions in space. Nomenclature A - area c - speed of sound CF - thrust coefficient cp - specific heat at constant pressure F - thrust g(Y) - eq. (4) / - impulse rate of mass flow Mach number velocity pressure

Verification Tests of Carbon-Carbon Composite Grids for Microwave Discharge Ion Thruster

Abstract: Anionbeam opticsfora10-cm-diam 400-W-classmicrowavedischargeion thrusterwasfabricatedanditsapplicabilityto along-termspacemissionwasdemonstrated.Theopticsconsistsofthree1-mm-thick e atcarbon ‐carbon composite panels with approximately 800 holes that were mechanically drilled and positioned with § 0:02-mmaccuracy.Whenmounted onanaluminum ring,spacingforthethreegridswaskeptat0.5 mm bythreesetsofspacers. The thruster produced an ion beam current of 140 mA with a microwave power of 32 W for plasma generation and a total acceleration voltageof 1.8 kV. Although thegrid is sputtered by the impingement of slow ions produced in charge-exchange collisions between fast beam ions and neutral atoms leaking from the engine, the grid showed only slight damage even after an 18,000-h endurance test. Also, other qualie cation tests including a mechanical test under launch conditions as well as a thermal vacuum test simulating the spacecraft thermal environment were successfully completed. Hence, the grid system was qualie ed for spacecraft propulsion.

Time- and Length Scales of Combustion in Liquid Rocket Thrust Chambers

Abstract: This paper summarizes experimental and computational results of research on propellant mixing and combustion in high-pressure LOX/GH2 rocket combustors. Hot-fire tests and numerical experiments are utilized to get more detailed information about the physics of mixing and combustion under high-pressure (transcritical) conditions. Time- and length scales of the reacting shear layer are identified by results from numerical simulation and experiments. The comparison between numerical and experimental results show the same tendencies of the distribution of length scales within the reactive shear layder. It was ascertained that the length scales are increasing downstream of the injector face plate. Qualitative and quantitative data about the flowfield (morphology, velocity, species distribution, etc.) at different boundary conditions are presented.

Numerical simulation of side loads in an ideal truncated nozzle

Abstract: The side loads induced by unsymmetrical and unsteady separation of the flow taking place in the nozzle extension during launch are a very important limiting factor for the performance of a rocket engine The onset of suchloads is not yet a fully understood phenomenon because only a limited amount of experimental data is available due to test difficulties A numerical study of the three-dimensional overexpanded nozzle flow during stationary operations has been undertaken to try to understand the origin of this phenomenon The flow separation in an truncated ideal contoured nozzle is investigated together with the resulting side loads Comparisons with experimental data are given The numerical simulation relies on the resolution of the three-dimensional unsteady Reynolds-averaged Navier-Stokes equations An algebraic turbulence model based on Baldwin-Lomax and Goldberg's backflow correction has been implemented in a three-zone formulation adapted to the flow topology of interest The main features of the flowfield, side-loads mean value, and the separation point location are well estimated With regard to the separation point location, the proposed method gives better results than the most commonly used semi-empirical criteria Unsteady characteristics of the flowfield are presented

Experimental and Analytical Design Verification of the Dual-Bell Concept

Abstract: The dual-bell nozzle concept has been investigated by means of analytical and experimental work. Based on earlier analytical and numerical work published by the authors, different dual-bell nozzles were designed and experimentally tested, in order to explore this concept regarding its aerodynamic characteristics. This experimental work included cold gas subscale tests and hot gas subscale tests, which were performed within a joint German/Russian Research Programme TEHORA, and complementary in the German National Technology Programme TEKAN. It is shown that depending on the type of nozzle contour used for the dual-bell nozzle extension, a sudden transition from sea-level to altitude mode operation can be achieved. Furthermore, important information from the hot gas tests regarding the wall heat transfer were gained.

Ethylene Ignition and Detonation Chemistry, Part 2: Ignition Histories and Reduced Mechanisms

Abstract: The detailed mechanism of 148 steps among 34 chemical species that has been described previously is reduced through steady-state and partial-equilibrium approximations to achieve simplified descriptions for use in ignition and detonation studies. The concentration histories for ethylene oxidation reveal different types of ignition processes under not very different conditions of pressure and temperature. When unimportant steps are sequentially deleted from the detailed mechanism, a short mechanism is developed involving only 38 irreversible elementary steps among 21 chemical species. This short mechanism is designed to be applied with reasonable accuracy over the entire range of conditions indicated previously. If attention is restricted to the induction period, then these 38 steps can be decreased to 21. During induction, the important chain carriers are mainly H, O, OH, and C 2 H 3 above 1500 K, whereas C 2 H 3 , HO 2 , and H 2 O 2 become dominant carriers below this temperature. At high temperatures, expressions for ignition time are derived by using radical depletion as the ignition criterion, whereas at lower temperatures, simplified thermal explosion theory in principle can be applied. Based on the detailed mechanism, a two-step mechanism involving only seven species is developed, which can facilitate computational investigations of multidimensional ignition and detonation processes of ethylene-air mixtures.