Showing papers in "Journal of Propulsion and Power in 1990"

Characteristics of multiple shock wave/turbulent boundary-layer interactions in rectangular ducts

Abstract: Multiple shock wave/turbulent boundary-layer interactions in a rectangular duct have been investigated using wall pressure measurements, surface oil flow visualization, spark schlieren photography, and laser Doppler velocimetry. Two undisturbed incoming Mach numbers were considered, Mach 2.45 and Mach 1.6. At Mach 2.45 the shock structure was a neutrally stable pattern of oblique shocks followed by repeated normal shocks with the level of flow confinement having only a small effect in the interaction. A large, three-dimensional separation region was observed. At Mach 1.6 the pattern consisted of a bifurcated normal shock followed by weaker, unbifurcated normal shocks. The boundary layer under the bifurcated shock was incipiently separated. In contrast to the Mach 2.45 case, the lower Mach number interaction was much steadier with the length of the interaction scaling directly with the level of flow confinement.

Stator/rotor interaction in a transonic turbine

Abstract: This paper presents calculations of a stator/rotor interaction in a highly loaded transonic first turbine stage. Of particular interest is the propagation and reflection of shocks which originate at the trailing edge of the upstream stator. These produce a 40% variation in the lift on the rotor, which would cause structural vibrations and increased losses. Also, the unsteady shocks would cause temporary boundary-layer separation near the leading edge of the suction surface. The numerical procedure solves the inviscid unsteady Euler equations, including quasi-three-dimensional terms. The use of a conservative treatment guarantees the correct treatment of the moving shocks. A simple technique is used to couple the calculations on the stator and rotor grids. A key feature of the paper is the use of "time-incline d" computational planes to allow the analysis of cases in which the ratio of stator and rotor pitches is not equal to unity or the ratio of two small integers.

Internal flow field studies in a simulated cylindrical port rocket chamber

Abstract: The objective of these studies is to experimentally characterize the mean and fluctuating flow field that develops along the length of a simulated cylindrical port rocket chamber. Flow simulation was accomplished by injecting ambient temperature nitrogen uniformly along the walls of 10.2-cm (4-in.) diam, porous-tube chambers connected to a choked sonic nozzle. Experiments were conducted with chamber L/D ratios of 9.5 and 14.3, at injection Mach numbers and Reynolds numbers typical of rocket motor values. Maximum Reynolds numbers based on injection and centerline velocities were, respectively, 1.8 x 10 and 1.6 x 10. Mean and fluctuating speed and turbulent shear stresses were measured in the principle coordinate directions using three-element hot-wire anemometers. The data show that noticeable velocity fluctuations in the head-end region generally decrease in intensity, relative to centerline speed, over the first five port diameters. At this point, regular velocity oscillations appear near the wall, just prior to the transition to turbulent flow. The oscillation frequency characteristics suggest the occurrence of vortical disturbances which exhibit pairing as they move away from the wall. The downstream turbulence development is characterized by a slow spreading toward the centerline: peak values of turbulence intensity and shear stress occur a few tenths of a port radius from the wall and remain relatively constant. Mean velocity profiles prior to transition show fair agreement with those derived for a rotational inviscid flow injected normal to the surface. A slow transition from these profiles occurs downstream in the turbulent region. Two surprising features of the flow were the occurrence of both buoyant flow influences and flow spinning in forward regions of the chamber.

Practical active control-system for combustion oscillations

Abstract: A low-frequency combustion instability of a flame burning in a duct has been successfully controlled by the unsteady addition of extra fuel. A suitably phased addition of only 3% more fuel reduces the peak in the pressure spectrum due to the combustion instability by some 12 dB. The acoustic energy in the 0-400 Hz bandwidth is reduced to 18% of its uncontrolled value. Since relatively little unsteady fuel is necessary, the mechanical power requirements of the controller are modest and the system is easy to implement.

Numerical simulations of an oblique detonation wave engine

Abstract: An account is given of the numerical methods employed in a code for the simulation of supersonic combustion, which is then applied to the simulation of attached detonations and flames associated with the oblique-detonation wave supersonic combustor concept. The addition of heat by a detonation wave results in a shorter combustor than can be obtained in more conventional scramjet designs. Pure oblique detonations have been produced in a stoichiometric, uniformly mixed hydrogen/air stream; the wave rotates upstream with energy release, according to simple analytical arguments. Flow visualization maps for Mach number and temperature are presented.

Navier-Stokes analyses of the redistribution of inlet temperature distortions in a turbine

Abstract: The flow exiting the combustor and entering the turbine of a gas turbine engine is known to contain both spatial and temporal variations in total temperature. Although historically it has been presumed that the turbine rotor responded to the average temperature, recent experimental evidence has demonstrated that the rotor actually separated the hotter and cooler streams of fluid so that the hotter fluid migrated toward the pressure surface and the cooler fluid migrated toward the suction surface. In the present study a time-accurate, two-dimensional, thin-layer, Navier-Stokes analysis of a turbine stage was used to analyze this phenomenon. The rough qualitative agreement between the measured and the computed results indicated that the analysis had successfully captured many of the important features of the flow.

Fundamental concepts of vectored propulsion

Abstract: Future fighter aircraft may maneuver, especially in the post-stall (PS) domain, by simultaneousl y directing their jets in the yaw, pitch, and roll coordinates. Consequently, thrust vectoring (TV) may gradually become a key element in helping fighters to survive and win in the close-combat arena. It also provides fighter aircraft with short-takeoff-and-landing (STOL) capabilities. This paper first defines the fundamental concepts associated with pure, or with partial TV powerplants. It then demonstrates that propulsion engineering should be expanded to include such unorthodox engine-design criteria as those of TV maneuverability and controllability. Second, the fundamental concepts of pure vectored propulsion are employed to design, construct, and laboratory test a new type of simultaneous roll-yaw-pitch TV system. Vectored remotely piloted vehicles (RPVs) were then constructed "around" these new propulsion systems. Flight tests of these RPVs since May 1987 have verified the STOL capability and enhanced maneuverability and controllability designable into vectored propulsion systems. They also became the first flight tests of pure vectored propulsion systems. The integrated methodology of laboratory/vectored-RPV-flight tests, as developed for this investigation, has been verified as cost effective and timesaving. Using this methodology a follow-up program was recently launched to help upgrade existing fighter aircraft, such as the F-15, F-16, and F-18, to become partially vectored PS aircraft. Finally, the basic conceptual changes associated with the very introduction of TV engines are summed up in terms of greater emphasis on highly integrated engine/flight-control testing methodologies and on reassessment of conventional concepts.

Mixing of an acoustically pulsed air jet with a confined crossflow

Abstract: Mean velocity and turbulence profiles, downstream of the jet orifice, in a 10 m/s crossflow have been measured over a range of Strouhal numbers and excitation powers for a jet/crossflow velocity ratio range of 1.3 to 4.6. This showed that acoustically exciting a jet flow causes significant increases in jet spread, penetration (up to 92% increase), and mixing. The jet mixing length was strongly reduced. Toroidal vortices were shown to be shedding from the jet orifice and produced profound changes in the jet structure. Increase of jet penetration and turbulence (hence mixing) began to saturate by about 80-W driving power, thus only small further gains were possible up to the maximum power used of 160 W. The jet turbulence and penetration data showed that the response appeared to be optimum at about a Strouhal number of 0.22. Overall, the jet mixing processes were significantly improved, in a controllable manner, by pulsating the jet flow.

Burning rate characteristics of GAP propellants

Abstract: The burning rate and flame structure of glycidyl azide polymer (GAP)-based composite propellants were examined in order to obtain a wide spectrum of burning rates. Crystalline fine particles of ammonium perchlorate, cyclotetramethylene tetranitramine, or triaminoguanidine nitrate, were mixed within GAP to formulate GAP propellants. Since GAP is an energetic self-sustaining combustible polymer, the burning rate characteristics of GAP propellants appeared to be fundamentally different from those of conventional composite propellants. Measured results indicate that the burning rate, pressure exponent, temperature sensitivity, and flame structure depend largely on the concentration of the crystalline additives. The observed burning rate characteristics were correlated with the concentration of the crystalline particles.

Numerical Investigation of Hot Streaks in Turbines

Abstract: A numerical method is developed for simulation of hot-streak redistribution in a two-dimensional model of a turbine rotor. The flow domain is divided into a viscous region near the blade where the Reynolds-averaged, thin shear-layer Navier-Stokes equations are solved using an implicit finite-volume technique, and an inviscid core region where the Euler equations are solved using an explicit finite-volume method. The computational mesh consists of an O-mesh and an H-mesh patched together smoothly to cover the domain of interest. Computations are performed using two different flow conditions. The first test case uses hot streaks with a temperature ratio of 1.2 and a tangential inflow angle of 40 deg, whereas the second test case is run with a temperature ratio of 2.0 and a tangential inflow angle of 45 deg. The computed solution from both test cases predicts a migration of hot gas to the pressure surface, which also has been observed experimentally. MODERN jet engine is designed to have an extremely high temperature gas leaving the combustor. The temper- ature tolerance of the guide vanes is usually based on an average value of the exit combustor temperature. Due to intro- duction of cooling air in the nozzle, the turbine entry temper- ature (TET) measured behind the guide vanes is lower than the combustor exit temperature. The temperature tolerance of the blades in the first rotor row is based on an averaged TET. Recent investigations, however, have shown that hot gas mi- grates to the pressure surface of the rotor blade. This can lead to peak temperatures on the rotor blade that might exceed acceptable metal temperatures and hence lead to blade fail- ures. This indicates that the TET, which is currently used, is too low an estimate for the rotor surface temperature. In an earlier work by Butler et al., 1 an experimental and analytical investigation of the redistribution process for an axial turbine stage was presented. In the experiment, a streak of hot air seeded with CO2 was introduced at one circumferen- tial location upstream of the inlet guide vane. The redistribu- tion of the hot streak was determined by measuring the con- centration of CO 2 inside the turbine stage. Measurements of CO2 taken on the rotor surface indicated that hot and cold gas had been segregated with the cold gas migrating to the suction surface and the hot gas to the pressure surface. In the same paper it was postulated that the segregation effect was due to the difference in rotor relative inlet angles of the hot and cold gases. The postulate is based on an observation by Kerrebrock and Mikolajczak2 in their work on wake transport in compres- sors. The observed segregation phenomenon1 has been investi- gated numerically by Rai and Dring.3 Their paper includes a comparison between Rai's computations and Butler's mea- surements. The computations did not give the same high tem- perature excess as was observed experimentally. The dis- crepancies between calculations and experiment were blamed on three basic differences in flow conditions: 1) the experi- ment was fully three-dimensional; 2) the temperature ratio between hot and cold gas was 1.2 in the calculations and 2.0 in

Status of xenon ion propulsion technology

Abstract: This paper describes a working-model xenon ion propulsion subsystem (XIPS) designed for north-south stationkeeping (NSSK) of 2500-kg-class geosynchronous communication satellites. The XIPS consists of a 25-cm-diameter laboratory-model thruster, a breadboard-model power supply, and a flight-prototype pressure regulator (the critical component of the pressure-regulated xenon feed system). With a thrust of 63.5 mN, specific impulse of 2800 sec, and thruster efficiency of 65 percent, the XIPS performance is believed to be the highest ever reported for an ion thruster operated at 1.3-kW input power. The XIPS power supply accepts an input power of about 1.4 kW from a 28- to 35-V bus and converts it into the seven outputs required for startup and operation of the thruster. The simplified power supply contains only about 500 parts and has demonstrated an unprecedented efficiency of 90 percent and a specific mass of about 8 kg/kW. The results of a highly successful wear-mechanism test in which the working-model XIPS was operated for 4350 hours and 3850 ON/OFF cycles are presented. These hours and cycles are equivalent to over ten years of NSSK on large communication satellites.

Multistep dump combustor design to reduce combustion instabilities

Abstract: Based on the understanding of the critical role of large-scale structures as drivers of pressure oscillations, a multistep dump was sucessfully tested to suppress pressure oscillations in a coaxial dump combustor. The multistep concept, which prevents development of large-scale structures, was studied in nonreacting air and water flows and in an annular diffusion flame before it was applied to the dump combustor burning gaseous fuel. The nonreacting tests in water and air showed that there is an optimal geometric configuration of the multisteps to achieve the highest level of turbulence. At this geometry the shear layer that is separated from the upstream step impinges on the next downstream step edge. When this geometry was tested in a dump combustor, the fuel injection pattern was found to be critical to obtain suppression of instabilities . With fuel injection distributed into the small-scale turbulence downstream of each step and not interfering with the flow impingement, the pressure oscillation in the dump combustor was suppressed.

Investigation of very high bypass ratio engines for subsonic transports

Abstract: A study of propulsion system performance for a family of high bypass ratio turbofan engines is presented. The bypass ratio range is from 6 to 17.5, and both bare engine performance and nacelle installation performance have been considered. Geared, variable pitch/variable nozzle engines with bypass ratios of 10.6, 14, and 17.5 have been studied parametrically, and the performance of a fixed pitch, gearless engine of bypass ratio 9.6 has been estimated to represent an advanced engine with conventional mechanical complexity. Nacelle performance data for a bypass 5 engine have also been included as a baseline. Thrust reverser design for these engines is also considered. Results of this study indicate that conventional nacelle installation losses do not reverse the trend of improving engine fuel efficiency with bypass ratio out to at least bypass ratio 17.5. If innovative concepts are used, such as reverse thrust from a reverse fan pitch and a short fan cowl, the installed fuel efficiency for highbypass-ratio engines looks even better. The bypass 9.6 engine shows a lower fuel burn benefit than the bypass 17.5 engine, but it offers the potential of reduced mechanical complexity and lower maintenance cost.

Thermal analysis of engine inlet anti-icing systems

Abstract: A hot air anti-icing system of a gas turbine engine inlet is analyzed numerically. A three-dimensional potential flow code, which accounts for compressibility effects, is used to determine the flowfield in and around the inlet. A particle trajectory code is developed using a local linearization technique. The trajectory code is used to calculate local water impingement rates. Energy balances are performed on both the surface runback water and the metallic skin to determine their temperature distributions. A variety of test cases are considered in order to validate the various numerical components of the process as well as to demonstrate the procedure.

Negative matter propulsion

Abstract: Negative matter is a hypothetical form of matter whose active-gravitational, passive-gravitational, inertial, and rest masses are opposite in sign to normal, positive matter. Negative matter is not antimatter, which as far as is known has normal (positive) mass. If an object made of negative matter could be obtained and coupled by elastic, gravitational, or electromagnetic forces to an object containing an equal amount of positive matter, the interactions between the two objects would result in an unlimited amount of unidirectiona l acceleration of the combination without the requirement for an energy source or reaction mass. In this paper, it is shown in exhaustive detail that, despite their unbelievable propulsive capabilities, negative matter propulsion systems do not violate the Newtonian laws of conservation of linear momentum and energy. Thus, logical objections to the existence of negative matter must be found elsewhere than in Newtonian mechanics. Suggestions are made where evidence for the existence of negative matter might be found.

Prediction of self-pressurization rate of cryogenic propellant tankage

Abstract: The SOLA-ECLIPSE code is being developed to enable prediction of the behavior of cryogenic propellants in spacecraft tankage. A brief description of the formulations used for modeling heat transfer and for determining the thermodynamic state is presented. Code performance is verified through comparison to experimental data for the self-pressurization of scale-model liquid hydrogen tanks. SOLA-ECLIPSE is used to examine the effect of initial subcooling of the liquid phase on the self-pressurization rate of an on-orbit full-scale liquid hydrogen tank typical for a chemical-propulsion orbit transfer vehicle. The computational predictions show that even small amounts of subcooling will significantly decrease the self-pressurization rate. Further, if the cooling is provided by a thermodynamic vent system, it is concluded that small levels of subcooling will maximize propellant conservation.

Modes of oscillation in a nonreacting ramjet combustor flow

Abstract: The results of the numerical simulation of cold flows in a ramjet are used to identify the mechanism that leads to the observed pressure fluctuations. The acoustic disturbance is defined as the unsteady part of the potential field and is shown to be driven by the instantaneous dilatation field. The quadrupole nature of the sound source around each vortex in the flowfield is demonstrated. The dilatation field in the vortex impingement region on the nozzle wall is considered a compact acoustic source and is analyzed by multipole expansion of the distributed field, revealing a strong axial acoustic dipole 180 deg out of phase with the impinging vorticity fluctuations. This dipole response to the vortical fluctuation is applied as the impedance for the vorticity/acoustic fluctuations at the nozzle. The spectral analysis of pressure and vorticity fluctuations reveals both a resonant acoustic mode in which the vortical disturbances excite the acoustic free modes, and a coupled mode in which the acoustic and vortical disturbances are coupled through dipole radiation at the nozzle and the acoustic susceptibility of the separating shear layer at the dump plane. A model for the coupled mode is proposed that provides a method for estimating its frequency.

Numerical simulations of oscillatory cold flows in an axisymmetric ramjet combustor

Abstract: A numerical simulation and analysis technique was developed to study the interaction between the vorticity component and the acoustic component of the flowfield in an axisymmetric ramjet combustor. To exclude the effects of the outflow boundary conditions, the interior of the combustor was isolated from the external region by a choked nozzle. The subsonic inflow boundary conditions are damping and thus do not force the oscillations. The shear layer separating at the step is perturbed by the low-frequency pressure fluctuation at the base of the step. The perturbation is then amplified downstream by the shear-layer instability and large-scale vortical structures are formed. When these structures impinge on the downstream nozzle wall, secondary vortices are generated, and large fluctuations in the Mach number upstream of the nozzle throat are observed. However, the Mach number in the supersonic region downstream of the throat remains stationary. The pressure and the vorticity spectra indicate the presence of the same low-frequency components, suggesting a possible vortex/acoustic wave interaction. Analysis of the computed mean-flow quantities and the higher moments indicates that, for the majority of the flow region, the large-scale structures are the main contributor to the transport of momentum. In a related study, the information obtained from these simulations is used to propose a model for one of the possible vortex/acoustic wave interactions.

Fusion as Electric Propulsion

Abstract: Fusion rocket engines are analyzed as electric propulsion systems, with propulsion thrustpowertoinputpower ratiothe thrustpower "gain," G�� much greater than unity. Gain values of conventionalsolar, fissionelectric propulsion systems are always quite smalle.g., G� < 0.8� . With these, "highthrust" interplanetary flight is not possible, because their system accelerationatcapabili� ties are always less than the local gravitational accelera� tion. In contrast, gain values 50� 100 times higher are found for some fusion concepts, which oer "high� thrust" flight capability. One performance example shows a 53.3 day � 34.4 powered; 18.9 coast� , oneway transit time with 19� payload for a singlestage Earth/ Mars vehicle. Another shows the potential for highac� celerationa� = 0.55 g0� flight in Earth/moon space.

Aerodynamic interference of vertical axis wind turbines

Abstract: The laminar flowfield and performance of clusters of two-dimensional vertical axis wind turbines are analyzed by idealizing the rotors as momentum sources. The flowfield dominated by the pressure field of the operating turbines is determined by solving the incompressible Navier-Stokes equations and mutual interference is observed to be elliptic in nature. Physical positioning of the turbines with respect to each other significantly affects the aerodynamic performance of the turbines. Several cases are examined starting with one pair of turbines to a cluster as big as 21 turbines. From the two-dimensional cases presented here, it is concluded that downwind turbine performance can be improved by judiciously choosing the angular relative orientation of the turbines. Qualitative analysis of the laminar flowfield is conducted with the help of streamlines and pressure contours. Performance comparisons of the turbines are made with a stand-alone turbine operating under similar flow conditions. This investigation lays the groundwork for the optimization of a wind farm.

Combustion of polymethylmethacrylate in a solid fuel ramjet

Abstract: The combustion behavior of polymethylmet hacrylate (PMMA) in a solid fuel ramjet (SFRJ) was investigated using a connected pipe test facility. The study was oriented towards understanding the most important phenomena related to the combustion in an SFRJ. Regression rates have been determined, and spectroscopy has shown the occurrence of OH, C2, and CH. Combustion pressure was shown to affect soot formation. No grain length effect on regression rate could be detected, whereas oxygen content in the air and inlet air temperature are important. Temperatures in the center of the bore are about 900 K, while temperatures in the recirculation zone approximate 1400 K. The highest temperatures were found in the near-wall region downstream of the recirculation zone. The combustion efficiency varied between 70% and 90%, and may be increased by increasing the fuel grain length and/or the oxygen content.

Multicomponent spray computations in a modified centerbody combustor

Abstract: The flow properties of the near-wake region of a ducted, bluff-body combustor are presented, based upon finite-difference computations involving turbulent, reacting flows with polydisperse and multicomponent sprays The analysis advances the state-of-the-art for the representation of the liquid-phase and for the coupling between the phases The gas phase representation is existing state-of-the-art Both droplet vaporization and turbulent mixing rates are found to be rate-controlling in the primary reaction region, and the mixing rate is identified to be the rate-controlling process in the later stages of combustion Under the inflow conditions leading to an equivalence ratio of one based on the primary air/fuel mass flow rates, the primary reaction region is found to be fuel-rich Enhanced energy conversion is observed to result from improved mixing; the effects of fuel composition are found to be negligible Near the wake-extinction limit of the combustor the effects of fuel composition are found to be very significant

Comparison of combustion experiments and theory in polyethylene solid fuel ramjets

Abstract: A research program was carried out to validate numerical simulation of the flow and combustion processes in the combustion chamber of a solid fuel ramjet with experimental results Experimental data were obtained by burning cylindrical fuel grains made of polyethylene in a solid fuel ramjet using a connected pipe facility For numerical simulation a computer code describing two-dimensional, steady-state turbulent flows through channels with and without a sudden expansion was developed Three different combustion models are incorporated, one based on finite-rate chemical kinetics, the other two based on the diffusion flame concept For the validation, emphasis was laid on comparing regression rate data in relation to chamber pressure, air mass flow, inlet air temperature, and step height Attention was also paid to reattachment length, temperature, and C2 and CH concentrations In some cases, a comparison with findings from other investigators was also made The results show good agreement between predicted and observed behavior downstream of the recirculation zone In the recirculation zone, however, the agreement is rather poor and can be attributed to the inability of the k-e model to predict heat transfer behind a rearward facing step accurately It is shown that the effective heat of gasification of the fuel is an important parameter A better understanding of its behavior in relation to combustion chamber conditions is needed Computed regression rate data are relatively insensitive to the combustion model employed

Modeling of a Small Hydrazine Thruster Plume in the Transition Flow Regime

Abstract: Three different techniques commonly employed in the calculation of rocket and thruster expansion plumes are assessed. These techniques vary both in computational expense and in the accuracy and detail of the solutions that they provide. The assessment is made with reference to the plume expanding from a small monopropellant hydrazine thruster and includes comparison with experimental data. Two of the modeling techniques, the Simons model and the Method of Characteristi cs, rely on the continuum equations. The third, the Direct Simulation Monte Carlo method (DSMC), adopts a discrete particle approach. The validity in employing continuum methods in the flowfield between the continuum and free molecular limits (i.e., the transition flow regime) is investigated. It is noted that the more computationally intensive DSMC solution method is the proper technique in this region of the expansion plume. Additional results provided solely by the DSMC calculations, such as thermal nonequilibrium effects, are presented. The consequences arising from the apparent differences in the results obtained with the continuum and discrete particle methods at the free molecular limit are assessed in terms of impingement effects.

Nonlinear finite-element analysis to predict fan-blade damage due tosoft-body impact

Abstract: A transient, material and geometric nonlinear, finite-element-based impact analysis, PW/WHAM, is presented. PW/WHAM couples the WHAM program, a transient geometric and material nonlinear plate finiteelement analysis, a fluid finite-element projectile, and contact algorithms to form an advanced numerical tool used to predict impact damage on structural components. The impacted component is modeled via WHAM plate finite elements while the impactor may be modeled with WHAM elements (hard-body impact) or fluid finite elements (soft-body impact). The direct integration time marching solution uses the explicit form of the Newmark B-Method difference equations. The spherically shaped fluid finite elements, utilizing a compressible nonlinear fluid constitutive law, are assembled in a closest-packed formation producing the soft-body impactor shape. General contact algorithms continualy track the bird/structure interaction eliminating the need for empirical tracking corrections. Several examples demonstrate the excellent correlation with test experience.

Observations of breakup processes of liquid jets using real-time X-ray radiography

Abstract: Liquid jet breakup mechanisms and processes have been studied extensively over the last one hundred years. However, since the region near the jet injector is too dense and optically opaque, conventional visualization cannot be applied with satisfactory experimental results. To unravel the liquid Jet breakup process in the non-dilute region, a newly developed system of real-time X-ray radiography together with an advanced digital image processor and a high-speed video camera was used in this study. Based upon recorded X-ray images, the inner structure of a liquid jet during breakup was observed. The Jet divergent angle, Jet breakup length, and void fraction distributions along the axial and transverse directions of liquid Jets, etc., were determined in the near-injector region. Both walland free-Jet tests were conducted to study the effect of wall friction on the Jet breakup process.

Experiments on a repetitively pulsed electrothermal thruster

Abstract: This paper presents experimental results from an investigation of a pulsed electrothermal (PET) thruster using water propellant. The PET thruster is operated on a calibrated thrust stand, and produces a thrust to power ratio of T/P = 0.07 + or - 0.01 N/kW. The discharge conditions are inferred from a numerical model which predicts pressure and temperature levels of 300-500 atm and 20,000 K, respectively. These values in turn correctly predict the measured values of impulse bit and discharge resistance. The inferred ideal exhaust velocity from these conditions is 17 km/sec, but the injection of water propellant produces a test tank background pressure of 10-20 Torr, which reduces the exhaust velocity to 14 km/sec. This value corresponds to a thrust efficiency of 54 + or - 7 percent when all experimental errors are taken into account.