Showing papers in "Journal of Propulsion and Power in 1989"
TL;DR: In this paper, the authors advocate a strategy for controlling turbomachine instabilities, whose primitive phases can be modeled by linear theory, but that eventually grow into a performance-limiting modification of the basic flow.
Abstract: In this paper, we advocate a strategy for controlling a class of turbomachine instabilities, whose primitive phases can be modeled by linear theory, but that eventually grow into a performance-limiting modification of the basic flow. The phenomena of rotating stall and surge are two very different practical examples in which small disturbances grow to magnitudes such that they limit machine performance. We develop a theory that shows how an additional disturbance, driven from real-time data measured within the turbomachine, can be generated so as to realize a device with characteristics fundamentally different than those of the machine without control. For the particular compressor analyzed, the control increases the stable operating range by 20% of the mean flow. We show that active control can also be used to destabilize a compressor in an undesirable state such as nonrecoverable stall. Examination of the energetics of the controlled system shows the required control power scales with the square of the ambient disturbance level, which can be several orders of magnitude below the power of the machine. Brief mention is also made of the use of structural dynamics, rather than active control, to enhance stability.
273 citations
TL;DR: The results of an experimental program aimed at determining the extent of the redistribution of an inlet temperature distortion in an axial flow turbine stage are presented in this paper, where air, seeded with CO2, was introduced at one circumferential location upstream of the inlet guide vane.
Abstract: The results of an experimental program aimed at determining the extent of the redistribution of an inlet temperature distortion in an axial flow turbine stage are presented. The program was conducted in a large-scale, low-speed, single-stage turbine where air, seeded with CO2, was introduced at one circumferential location upstream of the inlet guide vane. The migration of the seeded air through the turbine was determined by sensing CO2 concentration inside the stage. A temperature distortion was introduced by heating the seeded air. The CO2 concentration contours measured downstream of the vane showed h'ttle change with heating, indicating that the vane flowfield was relatively unaffected by the introduction of the temperature distortion. However, the CO2 contours observed on the rotor airfoil surfaces for the case with inlet heating indicated segregation of hot and cold gas with the higher temperature gas migrating to the pressure side and the lower temperature gas migrating to the suction side. Significant increases in rotor secondary flow were also observed. Two separate physical mechanisms are postulated to explain the observed experimental trends. Calculations performed by using a three-dimensional Euler solver show qualitative agreement with the experimental data.
192 citations
148 citations
TL;DR: In this paper, a finite-difference, unsteady, thin-layer Navier-Stokes approach to calculate the flow within an axial turbine stage is presented, where the relative motion between the stator and rotor airfoils is made possible with the use of patched grids that move relative to each other.
Abstract: Fluid flows within turbomachinery tend to be extremely complex. Understanding such flows is crucial to efforts to improve current turbomachinery designs, and the computational approach can be used to great advantage in this regard. This study presents a finite-difference, unsteady, thin-layer Navier-Stokes approach to calculating the flow within an axial turbine stage. The relative motion between the stator and rotor airfoils is made possible with the use of patched grids that move relative to each other. The calculation includes end-wall and tip-leakage effects. The numerical methodology is presented in detail in the present paper (Part I). The computed results and comparisons of these results with experimental data are presented in a companion paper (Part II).
122 citations
TL;DR: A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solid propellant rocket motor chambers.
Abstract: A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier-Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solid propellant rocket motor chambers. The calculation procedure incorporates a two-equation (k-epsilon) turbulence model and utilizes a consistently split, linearized block-implicit algorithm for numerical solution of the governing equations. The code was validated by comparing computed results with the experimental data obtained in cylindrical-port cold-flow tests. The agreement between the computed and experimentally measured mean axial velocities is excellent. The axial location of transition to turbulent flow predicted by the two-equation (k-epsilon) turbulence model used in the computations also agrees well with the experimental data. Computations performed to simulate the axisymmetric flowfield in the vicinity of the aft field joint in the Space Shuttle solid rocket motor using 14,725 grid points show the presence of a region of reversed axial flow near the downstream edge of the slot.
114 citations
TL;DR: Wood et al. as discussed by the authors used interferometric measurements of droplet size and droplet velocity to establish a fuel of controlled composition for modeling, and for the study of fuel property and chemical composition effects in the combustion of JP-4 fuels.
Abstract: Author(s): Wood, CP; Mc donell, VG; Smith, RA; Samuelsen, GS | Abstract: A surrogate fuel comprised of 14 pure hydrocarbons is formulated based on the distillation curve and compound class composition of a petroleum-derived JP-4. The goal is to establish a fuel of controlled composition for modeling, and for the study of fuel property and chemical composition effects in the combustion of JP-4 fuels. Spatially resolved interferometric measurements of droplet size and droplet velocity are obtained and compared for both the petroleum and surrogate JP-4 in a nonreacting spray chamber. Measurements are also obtained for a high aromatic JP-5 of purposefully disparate properties. The performance of these three fuels is then compared in a swirl-stabilized, spray-atomized model laboratory combustor where in-flame measurements of velocity and temperature are acquired and compared. The nonreacting measurements of atomization quality establish that the atomization characteristics of the petroleum and surrogate JP-4 are identical, whereas the atomization performance of the JP-5 is significantly different. Under reacting conditions, substantial differences between the JP-4 and JP-5 are observed in both the velocity and thermal fields, whereas the surrogate, in contrast, yields an identical velocity and thermal field to that of the petroleum JP-4. © 1989, American Institute of Aeronautics and Astronautics, Inc., All rights reserved.
95 citations
TL;DR: In this paper, the mixing problems in hypervelocity scramjet combustors are discussed, and techniques for providing turbulence and/or mixing enhancement are described numerically and options for producing oscillatory shock waves for mixing augmentation are discussed.
Abstract: The paper discusses mixing problems in hypervelocity scramjet combustors. Techniques for providing turbulence and/or mixing enhancement are described. One such technique, the oscillating shock interaction, is studied numerically and options for producing oscillatory shock waves for mixing augmentation in scramjet combustors are discussed.
92 citations
TL;DR: A small uncooled plasma torch was developed and used in combination with an injector designed to study ignition and flameholding in hydrogen-fueled supersonic flows as discussed by the authors.
Abstract: A small, uncooled plasma torch was developed and used in combination with an injector designed to study ignition and flameholding in hydrogen-fueled supersonic flows. The plasma torch was operated on mixtures of hydrogen and argon with total flows of 10 to 70 scfh. The fuel injector design consisted of five small upstream pilot fuel injectors, a rearward facing step for recirculation, and three main fuel injectors downstream of the step. The plasma torch was located in the recirculation region, and all injection was perpendicular to the Mach 2 stream. Both semi-freejet and ducted tests were conducted. The experimental results indicate that a low power plasma torch operating on a 1:1 volumetric mixture of hydrogen and argon and located in the recirculation zone fueled by the upstream pilot fuel injectors is a good igniter for flow conditions simulating a flight Mach number of 3.7. The total temperature required to autoignite the hydrogen fuel for this injector geometry was 2640 R. The injector configuration was shown to be a good flameholder over a wide range of total temperature. Spectroscopic measurements were used to verify the presence of air total temperatures below 1610 R.
84 citations
TL;DR: In this paper, the mixing characteristics of circular, small-aspect-ratio elliptic and rectangular jets were studied in subsonic, sonic, and supersonic flows.
Abstract: The mixing characteristics of circular, small-aspect-ratio elliptic and rectangular jets were studied in subsonic, sonic, and supersonic flows. The experiments were carried out in both cold and hot flows using hot-wire anemometry, thermocouples, and photography. The elliptic and rectangular jets had similar features, with a slightly better mixing performance of the elliptic jet in the subsonic and supersonic flows. The elliptic and rectangular jets had a higher spreading rate relative to the circular jet, epecially at the minor axis plane. In the subsonic jet, the spreading rate was limited to the first five equivalent diameters (De) from the nozzle. In the supersonic underexpanded jet, the spreading rate was three times higher in the entire range (30£>e) measured. The minor axis plane was also characterized by high intensity of near-field pressure fluctuations. The two phenomena can be related to each other when an acoustic feedback occurs. Nomenclature & = aspect ratio D = circular nozzle diameter De = equivalent diameter E = energy of the fluctuating pressure components in the power spectrum / = frequency M = Mach number r = radial coordinate Re = Reynolds number, U0De/v R05 = half-width of the jet; r at with U= U^ /2 Rw 5 = half-width of the reacting jet; r at which
74 citations
TL;DR: In this paper, the geometric and size effects on the cumbustion of solid fuel ramjets were investigated, and the results indicated that the local regression rate is closely related to the local convective heat flux.
Abstract: The paper summarizes an experimental investigation concerning the geometric and size effects on the cumbustion in solid fuel ramjets (SFRJs). Polymethylmethacrylate (PMMA) solid fuel was used, and the combustor simulated conditions resulting from flight at sea level and Mach 3, Instantaneous and local fuel regression measurements indicated the following conclusions: the local regression rate is closely related to the local convective heat flux; the honuniformity of fuel regression rate has an attenuating effect on the dependence of the mean regression rate on the mass flux in extended burn-time tests; the regression pattern is not affected by downstream conditions; and mean regression rate decreases with increasing port diameter. Nondimensional scales normalized by the port diameter were found to give generalized expressions for different motors.
61 citations
TL;DR: In this article, three oblique detonation ramjet-in-tube drive modes are presented; the operational velocities for the present results range from 3.5 to 10.0 km/s.
Abstract: Ramjet-in-tube techniques have been proposed to accelerate masses up to thousands of kilograms to velocities of 0.7-12.0 km/s by chemical means. CFD calculations for three oblique detonation ramjet-in-tube drive modes are presented; the operational velocities for the present results range from 3.5 to 10.0 km/s. The first drive mode achieves ignition on the reflection of the nose cone bow shock. The second drive mode relies on a sudden, steep, but small increase in projectile radius to initiate a detonation wave, following a deliberately gentle, gradual compression process. The third drive mode is similar to the second mode except that the projectile is thermally protected by flying it through a core of pure hydrogen gas surrounded by a detonable mixture. At optimum operating conditions, the thrust pressure ratios (defined as the effective thrust pressure divided by the maximum cycle pressure) for the three modes range from 0.12 to 0.30, and the efficiencies (defined as the thrust times the velocity divided by the rate of chemical energy release) range from 0.09 to 0.26. Tables of thrust pressure ratios and efficiency data and representative plots of the pressure fields around the projectiles are presented.
TL;DR: In this paper, the effects of oscillating gas pressure and velocity on droplet vaporization rates during combustion instability were examined, and the gain or response function associated with the oscillatory vaporization rate component in phase with the pressure was shown to be sufficiently large to sustain instability.
Abstract: The effects of oscillating gas pressure and velocity on droplet vaporization rates during combustion instability are examined. Droplets are continuously injected through the pressure cycle, and the behavior of each droplet is calculated. Individual droplet vaporization rates are combined to give the total vaporization rate for a spray. Heat conduction and convection within the droplet interior are considered and found to have very significant effects compared to the infinite-conductivity or rapid-mixing models of the past. The gain or response function associated with the oscillatory vaporization rate component in phase with the pressure is shown to be sufficiently large to sustain instability.
TL;DR: In this article, an aerothermochemical analysis for the process of carbon-carbon composite material regression in large advanced solid-propellant rocket motors has been conducted, with the main idea of the nozzle regression being due to the carbon chemical attack by H2O.
Abstract: An aerothermochemical analysis for the process of carbon-carbon composite material regression in large advanced solid-propellant rocket motors has been conducted. The analytical approach is similar in spirit to the approach of Klager, Keswani, and Kuo, with the main idea of the nozzle regression being due to the carbon chemical attack by H2O. The different steps of the work have consisted of the development and applications of several numerical codes substantiated by experimental results concerning the regression rate and the surface roughness of a carbon-carbon material. The calculated results show good agreement between measured data and the predicted regression when a flow transition is assumed, in the model, between a laminar boundary layer existing on the "smooth" virgin carbon-carbon material at the firing start and a turbulent boundary layer existing on the very rough ablative carbon-carbon surface during stabilized motor operation.
TL;DR: In this article, an inviscid, two-dimensional numerical scheme coupled to the detailed reaction kinetics of combustion was developed for the oblique detonation wave engine, which couples the chemistry via an operator-splitting method.
Abstract: The concept of oblique detonation waves has been suggested as an efficient mechanism for ignition of the combustible mixture in a scramjet. To verify this concept, both experimental and numerical capabilities must be developed. We have developed an inviscid, two-dimensional numerical scheme coupled to the detailed reaction kinetics of combustion. The scheme is second-order-accurate total variation diminishing, time accurate, and couples the chemistry via an operator-splitting method. The code is fully vectorized for maximal efficiency and has been applied to the study of oblique detonations at various conditions of pressure and Mach number. The method is discussed in the paper, as well as the numerical results pertinent to the proof of concept of the oblique detonation wave engine.
TL;DR: In this article, a finite-difference, unsteady, thin-layer Navier-Stokes solution to the flow within an axial turbine stage is presented, which includes end-wall and tip-leakage effects.
Abstract: Fluid flows within turbomachinery tend to be extremely complex. Understanding such flows is crucial to efforts to improve current turbomachinery designs, and the computational approach can be used to great advantage in this regard. This study presents a finite-difference, unsteady, thin-layer Navier-Stokes solution to the flow within an axial turbine stage. The computational methodology developed for this simulation is presented in Part I of this paper. The calculation includes end-wall and tip-leakage effects. Results in the form of time-averaged surface pressures, pressure amplitudes (corresponding to the pressure fluctuation in time), near-surface velocity vectors, and pressure contours in the passage areas are presented. The numerical results are compared with experimental data wherever possible and the agreement between the two is found to be good.
TL;DR: In this article, a small solid-fuel ramjet (SFRJ) combustor was tested at a Mach number of 3 at sea level using a static test system with a 25kW electrical air heater.
Abstract: Experimental and analytical investigations of a small solid-fuel ramjet (SFRJ) combustor were conducted. A static test system with a 25-kW electrical air heater simulated the air temperature and pressure encountered in flight at a Mach number of 3 at sea level. The transparent polymethylmethacrylate fuel used in the tests permitted continuous video photography, revealing the local fuel-regression-rate behavior and the instantaneous ignition and combustion phenomena. The results demonstrated high combustion efficiency and indicated peculiar local and average fuel-regression-rate correlations. The analysis indicated that the specific conditions resulting from the low Reynolds number range in small SFRJ motors, in contrast to large combustors, enhance the effect of the sudden-expansion heat-transfer regime relative to the boundary-layer regime. 14 refs.
TL;DR: In this paper, the authors present some recent concepts in Mars Sample Return (MSR) missions that utilize extraterrestrial resources, including the power and energy needs of this mission, and the dramatic savings in Shuttle (or other) vehicle launches are quantitatively plotted.
Abstract: This paper presents some recent concepts in Mars Sample Return (MSR) missions that utilize extraterrestrial resources. The concepts examined include the power and energy needs of this mission. It is shown that solar energy is not especially attractive. Radioisotopic power generator and a Rankine cycle use are seen to be viable options. Quantitative estimates, taking into consideration state-of-the-art and projected technologies indicate that the power/energy per se is not critical to the mission - but reliability is. Hence, various modern options for the components of the power generation and utilization are discussed. The dramatic savings in Shuttle (or other) vehicle launches are quantitatively plotted. The basic system that is discussed here is the production of hydrocarbon (methane) fuel and oxygen from Martian atmosphere. For the simplest mission, it is seen that earth-carried methane burned with oxygen produced on site provides the best system.
TL;DR: In this paper, a computational fluid dynamics (CFD) code was developed to compute the mixing and combustion of hydrogen fuel in the turbulent flowfields of supersonic combustion ramjets (scramjet) with primary interest for the case of transversely injected fuel.
Abstract: A computational fluid dynamics (CFD) code has been developed to compute the mixing and combustion of hydrogen fuel in the turbulent flowfields of supersonic combustion ramjets (scramjet) with primary interest for the case of transversely injected fuel. The code solves the three-dimensional time-dependent Reynolds-averaged Navier-Stokes equations, including species transport for a four-species, two-reaction, global finite-rate-chemistry model. Numerical integration was obtained by MacCormack's explicit method, and turbulence was modeled by the eddy viscosity model of Baldwin-Lomax. Mass diffusion was based on Pick's Law, and unit Lewis number was assumed. The code has been validated by comparing computational results with existing experimental data and applied to simulate a realistic scramjet combustor flowfield, particularly in the vicinity of the fuel injectors. Results from the comparisons are presented and discussed.
TL;DR: The microwave electrothermal thruster as discussed by the authors is a special case of the microwave thermal thrusters, where no electrodes are in contact with the working fluid, since nonthermal, radiative mechanisms transfer the energy into the working fluids.
Abstract: The microwave electrothermal thruster shows promise for spacecraft propulsion and maneuvering. It promises advantages over other electrothermal thrusters in the areas of operating life, efficiency, and propellant selection. In the microwave thermal thruster, the electric power is first converted to microwave-fre quency radiation. In a specially designed microwave cavity system, the electromagnetic energy of the radiation is transferred to the electrons in a plasma sustained in the working fluid. The resulting high-energy electrons transfer their energy to the atoms and molecules of the working fluid by collisions. The heated working fluid expands through a nozzle to generate thrust. In the microwave electrothermal thruster, no electrodes are in contact with the working fluid, since nonthermal, radiative mechanisms transfer the energy into the working fluid. The main requirements for the materials of construction are that the walls of the discharge chamber be insulating and, at least in part, transparent to microwave radiation at operating conditions. Several experimental configurations of microwave electrothermal thrusters are described and compared. Diagnostic methods used to study microwave plasmas under conditions used in the thruster are described and selected results presented for titration, spectroscopy, calorimetry, electric field measurements, and gas-dynamic methods. Estimated performance efficiencies are reported and compared with other electrothermal systems. Results of computer modeling of the plasma and of the gas flowing from the plasma are summarized.
TL;DR: In this paper, time-dependent, compressible numerical simulations have been performed to study the flow fields in an idealized ramjet consisting of an axisymmetric inlet and a combustor.
Abstract: Time-dependent, compressible numerical simulations have been performed to study the flowfields in an idealized ramjet consisting of an axisymmetric inlet and a combustor. The simulations indicate strong coupling between the flowfield and the acoustics of both the inlet and the combustor. For the cases studied, forcing at the first longitudinal acoustic mode of the combustor induces vortex rollup near the entrance to the combustor at the forcing frequency. A low-frequency oscillation is also observed in all of the simulations. Pressure fluctuations in the inlet indicate that the low frequency corresponds to a quarter-wave mode in the inlet. Changing the length of the inlet appropriately changes the observed low frequency. The merging pattern of the vortices in the combustor is modified significantly when either the acoustics of the combustor or that of the inlet are changed. These merging patterns are explained on the basis of an interaction between the vortex-rollup frequency and the acoustic modes of the inlet and combustor.
TL;DR: In this paper, an integrated Xenon ion propulsion subsystem was designed to perform 10 years of stationkeeping for 2500-kg-class geosynchronous communications satellites, which achieved an efficiency of 65% using a highly efficient ring-cusp discharge chamber equipped with a three-grid ion extraction assembly and a simplified power supply.
Abstract: This paper describes an integrated xenon ion propulsion subsystem designed to perform 10 years of northsouth stationkeeping for 2500-kg-class geosynchronous communications satellites. The propulsion subsystem comprises a 25-cm-diam thruster, a power supply, and a propellant tankage and control unit. With a spacecraft bus power of 1400 W, the unit produces 63 mN of thrust at a specific impulse of 2900 s. A propulsion subsystem efficiency of 65% is obtained using a highly efficient ring-cusp discharge chamber equipped with a three-grid ion extraction assembly (thruster efficiency of 70%) and a simplified power supply that operates at 90% efficiency. Results show that the propulsion subsystems can be turned on with a single analog command, reaching its fullthrust level in less than 3 min.
TL;DR: In this paper, two theoretical models for calculating the current and flow distributions in self-field MPD thrusters have been developed and are applied to evaluate the effects of geometry, propellant type, scaling, and other parameters on the thruster performance.
Abstract: Two theoretical models for calculating the current and flow distributions in self-field MPD thrusters have been developed and are applied to evaluate the effects of geometry, propellant type, scaling, and other parameters on the thruster performance. For continuous thrusters, a stationary code has been developed. The extended Ohm's law is used to calculate the current contour lines, and a one-dimensional, two-component expansion flow model is applied to obtain the velocity, temperature, and pressure distributions for calculating the gas properties, which are again used in Ohm's law. An integration over the volume and thermal forces equals the thrust. The differential equation is solved by means of a finite-difference method for the geometry of the nozzle-type plasma thruster DT2-IRS, which has been investigated experimentally in a steady-state as well as in a quasi-steady-state mode. The calculated current density distribution and the computed thrust are compared with these experimental results. For the starting phase of the steady-state MPD thrusters as well as for pulsed thrusters, a time-dependent, fully two-dimensional code has been developed. It uses a modified McCormack FD method in cylindrical coordinates to calculate the time-dependent flow, temperature, and pressure fields.
TL;DR: In this paper, a theoretical model has been developed to investigate turbulent mixing and combustion processes in the main combustion chamber of a solid-propellant ducted rocket, based on Favre-averaged conservation equations with a two-step chemical reaction scheme and is solved by a semi-implicit finite-difference method.
Abstract: A theoretical model has been developed to investigate turbulent mixing and combustion processes in the main combustion chamber of a solid-propellant ducted rocket. The formulation is based on Favre-averaged conservation equations with a two-step chemical reaction scheme and is solved by a semi-implicit finite-difference method. Turbulence closure is achieved using a well-known k-e two-equation model. Calculated flow structures show good agreement with preliminary experimental results obtained from the schlieren flow-visualization study. The influences of various parameters, including dome height and inlet flow angle, on the propulsive performance of the system are investigated in detail.
TL;DR: In this article, the authors summarized communications and dynamic electromagnetic experiences using electromagnetic, electrostatic, and electrothermal propulsion systems, with particular attention paid to the performance of spacecraft subsystems and payloads during propulsion operations.
Abstract: Introduction A electric propulsion systems become ready to integrate with spacecraft systems, the characterization of propulsion system radiated emissions is of significant interest. This paper briefly summarizes communications and dynamic electromagnetic experiences using electromagnetic, electrostatic, and electrothermal propulsion systems. Electromagnetic radiated emission results from ground tests and flight experiences are presented, with particular attention paid to the performance of spacecraft subsystems and payloads during thruster operations. The impacts to transmission of radio frequency signals through plasma plumes are also reviewed. Over the last 30 years, more than 60 spacecraft were flown using electric propulsion systems for drag makeup, stationkeeping, and experiments." The major flight-qualified thruster systems are ablative pulsed plasma thrusters (PPT), ion thrusters, magnetoplasmadynamic (MPD) thrusters, and resistojets. Average power dedicated to the propulsion system ranged from about 3 W for early pulsed plasma devices to about 1 kW for each (space electric rocket test) SERT-II ion thruster." (See Table 1.) Table 1 does not characterize all electric propulsion flights; only publications that refer to thrusters electromagnetic (EM) emissions or RF interactions with plumes were considered in this survey. Except for some resistojet systems, most electric propulsion systems require power processing equipment to tailor the battery or solar array power to that of the thruster. The thruster/power processor produces an electromagnetic environment that could potentially impact spacecraft systems such as communications, guidance, navigation and control, payloads, and experiments. The EM environments may have permanent and varying magnetic fields along with radio frequency and conducted electrical emissions. Prior to flight, electromagnetic interference (EMI) measurements must be made, and hardware must be immunized or the level of EM I reduced to satisfy the compatibility requirements. Previous papers have surveyed the particle and field interactions using ion thrusters and MPD arcjets." This paper will review radio frequency interference (RFI) component test specifications, results of spacecraft integration tests, and radiated emissions from flight systems using electromagnetic, ion, and electrothermal thrusters. Spacecraft schematics and ground test configurations will be discussed, general test procedures will be synopsized, and spacecraft integration test results will be reported. Most of the EM emission results were obtained from four flights using ablative pulsed plasma thrusters, six ion propulsion flights, and three flights using electromagnetic devices (see Table 1). Only published or publicly available information was used in compiling this report. Most of the electric propulsion work pertaining to EM emissions characterization was performed in the U.S., Japan, USSR, West Germany, and China.
TL;DR: In this article, a systematic investigation of the effects on compressor performance of recessing (trenching) the case over the rotor tips was conducted in a four-stage research compressor.
Abstract: A systematic investigation of the effects on compressor performance of recessing (trenching) the case over the rotor tips was conducted in a four-stage research compressor. Variations in the slope and depth of the trench, rotor tip penetration into the trench, distances of start and termination of the trench relative to the blade edges, and rotor-tip clearances were evaluated. The effects of each geometric parameter on compressor efficiency and stall margin were determined. A new sloped trench configuration was developed and found to be superior to any other trench configuration tested, giving in one case a 0.5 point improvement in efficiency without adversely affecting stall margin.
TL;DR: In this article, a low-power, uncooled hydrogen plasma torch has been built and tested to evaluate its potential as a possible flame holder for supersonic combustion, and the stability limits of the torch are delineated and its electrical and thermal behavior documented.
Abstract: The residence time of the combustible mixture in the combustion chamber of a scramjet engine is much less than the time normally required for complete combustion. Hydrogen and hydrocarbon fuels require an ignition source under conditions typically found in a scramjet combustor. Analytical studies indicate that the presence of hydrogen atoms should greatly reduce the ignition delay in this environment. Because hydrogen plasmas are prolific sources of hydrogen atoms, a low-power, uncooled hydrogen plasma torch has been built and tested to evaluate its potential as a possible flame holder for supersonic combustion. The torch was found to be unstable when operated on pure hydrogen; however, stable operation could be obtained by using argon as a body gas and mixing in the desired amount of hydrogen. The stability limits of the torch are delineated and its electrical and thermal behavior documented. An average torch thermal efficiency of around 88 percent is demonstrated.
TL;DR: In this paper, an extensive numerical experiment has been conducted to evaluate rocket thruster performance using a laser-sustained hydrogen plasma as the propellant, which was sustained using a 30 kW CO2 laser beam operated at 10.6 microns focused inside the thruster.
Abstract: An extensive numerical experiment has been conducted to evaluate rocket thruster performance using a laser-sustained hydrogen plasma as the propellant. The plasma was sustained using a 30 kW CO2 laser beam operated at 10.6 microns focused inside the thruster. The steady-state Navier-Stokes equations coupled with the laser power absorption process have been solved numerically. A pressure based Navier-Stokes solver using body-fitted coordinate was used to calculate the laser-supported rocket flow which included both recirculating and transonic flow regions. The local thermodynamic equilibrium (LTE) assumption was used for the plasma thermophysical and optical properties. Geometric ray tracing was adopted to describe the laser beam. Several different throat size thrusters operated at 150 and 300 kPa chamber stagnation pressure were studied. It was found that the thruster performance (vacuum specific impulse) was highly dependent on the operating conditions, and a properly designed laser supported thruster can attain a specific impulse around 1500 secs. The heat loading on the thruster wall was also estimated and was in the range of that for a conventional chemical rocket.