Showing papers in "Journal of Propulsion and Power in 1992"

Historical development of glycidyl azide polymer

Abstract: The glycidyl azide polymer (GAP) is an energetic, thermally stable, insensitive, hydroxyl-terminated prepolymer with potential applications in advanced solid propellants including 1) class 1.3 nondetonable minimum smoke propellants, 2) clean propellants for solid rocket boosters, 3) gas generators/aircraft starter cartridges, 4) low-cost ASAT maneuver propulsion systems, and 5) high-performance space propellants for orbital transfer vehicles. The historical development of GAP from its conception is reviewed in this paper.

Unsteady analysis of hot streak migration in a turbine stage

Abstract: Two- and three-dimensional Navier-Stokes analyses are used to predict unsteady viscous rotor-stator interacting flow in the presence of a combustor hot streak. Predicted results are presented for a two-dimensional three-stator/four-rotor, a two-dimensional one-stator/one-rotor, and a three-dimensional one-stator/one-rotor simulation of hot streak migration through a turbine stage. Comparison of these results with experimental data demonstrates the capability of the three-dimensional procedure to capture most of the flow physics associated with hot streak migration including the effects of combustor hot streaks on turbine rotor surface temperatures. It is noted that blade count ratio has little effect on predicted time-averaged surface pressure and temperature distributions, but a substantial effect on the unsteady flow characteristics. It is shown that high-temperature hot streak fluid accumulates on the pressure surface of the rotor blades, resulting in a high time-averaged surface temperature 'hot spots'.

Hydrocarbon-fueled scramjet combustor investigation

Abstract: An experimental program is being conducted for developing the technology required by a hydrocarbon-fueled engine capable of operating as a scramjet in the Mach 5.6-7.0 range. A series of tests have been run to define scramjet combustor design criteria applicable to a scramjet configuration involving a pair of modular, axisymmetric combustors mounted on the lower surface of a supersonic/hypersonic missile. The airbreathing pilot combustor incorporates an external mainstream fuel injector which serves as the primary fuel injection stage for the supersonic combustor; it has been demonstrated that the pilot promotes the efficient combustion of either gaseous ethylene or preheated JP-5, when they are injected into the mainstream flow as a primary fuel.

Active control of unsteady combustion-induced oscillations

Abstract: A loudspeaker-based active control system has been used to study a longitudinal combustion instability mode and its active control in a small scale premixed combustor. A simple gain/phase-shift type controller is shown to successfully suppress oscillations in the combustor at all equivalence ratios at low flow rates. It is found, however, that as the flow rate through the combustor is increased, the performance of the controller deteriorates rapidly, both in terms of allowable gain and phase margin and reduced effect on dominant instability such that beyond certain flow rates the controller is unable to suppress the oscillations for any values of gain and phaseshift. Nyquist stability diagram obtained from the frequency response data are used to understand the performance of the controller and help design an advanced controller with somewhat improved performance. The physical model developed by Lang et al. is adapted to help understand important features of the combustion instability and its active control. Without the controller the model is found to predict the frequency of the dominant instability mode and its variation with equivalence ratio fairly well. For off-stoichiometric operation the model predicts an increase in the number of potential instability modes as is also observed in the measured power spectra. With the controller "on," the model shows that the controller interacts significantly with the combustion and in one case examined here, it is shown that the controller stabilizes the second longitudinal mode for one value of gain but at just a slightly higher value renders the previously stable third mode unstable. This offers plausible physical explanation for the observed limited gain and phase margin and the limited effectiveness of the controller at the higher flow rates. The model also predicts the stability boundary to have a reduced phase margin as observed experimentally.

Planar fluorescence imaging of a transverse jet in a supersonic crossflow

Abstract: Planar laser-induced fluorescence (PLIF) imaging has been used to examine the mixing and combustion of a sonic jet of gas injected transversely into a supersonic freestream flowing within a shock tube. Single-shot and frame-averaged fluorescence images have been acquired for nonreacting and reacting flows in side-view and end-view orientations. In the nonreacting experiments, nitric oxide seeded within the jet fluid was used to examine the penetration and mixing of the jet with the freestream, without the influence of chemical reaction and heat release. In the reacting experiments, the hydroxyl radical (OH), formed by the combustion of a hydrogen jet injected into an oxidizing freestream, was used to locate the reaction zones. The OH images indicate that combustion takes place primarily in the shear layer formed by the jet and the freestream, and in the boundary layer adjacent to the wall. For both the nonreacting and reacting results, the single-shot images show the presence of large-scale turbulent structures not apparent in the frame-averaged images. These results demonstrate the importance of examining and understanding the instantaneous flowfield, because it is the instantaneous, rather than mean state, of the flow that ultimately determines the extent to combustion.

Active control of nonlinear pressure oscillations in combustion chambers

Abstract: A theoretical analysis of active control of nonlinear acoustic instabilities in combustion chambers is developed in this article. The formulation starts with a generalized wave equation that describes the dynamic behavior of second-order nonlinear oscillations with distributed feedback actions. Control inputs are provided by the burning of the injected seconday fuel in the chamber, with its instantaneous mass flow rate modulated by a proportionalplus-integral (PI) controller located between the pressure sensor and the fuel injection mechanism. Various nonlinear stability characteristics, including the existence and stability of limit cycles, are studied analytically using the method of time averaging. In addition, an optimization procedure is developed for selecting controller gains. Nomenclature Anij = parameters associated with nonlinear acoustic coupling a = speed of sound in mixture Bnij = parameters associated with nonlinear acoustic coupling b = spatial distribution of burning of control fuel, Eq. (5b), Cv - constant-volume specific heat for two-phase mixture c = amplification factor of sensor output Dni = linear parameters, Eq. (11) Eni = linear parameters, Eq. (11) e - error signal between desired response and actual measurement

Supersonic nozzle mixer ejector

Abstract: The intent of this article is to describe recent experimental findings relative to supersonic nozzle mixer ejector performance. Such ejectors are a candidate means to reduce jet noise of commercial supersonic aircraft during takeoff and landing. The mixer ejector concept involves the introduction of an array of large-scale, low-intensity streamwise vortices into the downstream mixing duct, which enhances mixing through an inviscid stirring process. This results in increased ejector pumping performance and more completely mixed flows exiting the ejector shroud. Past experimental and analytical investigations of mixer ejectors have been confined to low-speed subsonic flows, and low primary temperatures (less than 2000°F). In this flow regime, ejector static pumping benefits of over 100% were achieved relative to conventional ejector designs. The goal of the present study was to evaluate mixer ejector performance in the high-temperature, supersonic primary flow regime. A convergentdivergent primary lobed nozzle (i.e., mixer nozzle) was designed and tested at choked pressure ratios in an ejector. Ejector pumping and exit plane mixing were measured for the mixer ejector and a conventional slot nozzle ejector. The two configurations were operated at a nozzle exit Mach number of 1.5 (nozzle pressure ratio = 3.4), a primary fluid total temperature of 1000°F, and a simulated forward flight Mach number of 0.1. Results indicate that properly designed lobed nozzles can increase supersonic ejector pumping by over 15%, relative to conventional slot primary nozzles.

High current hollow cathode phenomena

Abstract: Experimental results show that the energies of ions produced near a hollow cathode orifice can be several times the anode-to-cathode potential difference generally considered available to accelerate them. These energies (of order 50 eV) are sufficient to induce substantial sputter erosion rates. Increases in discharge current (to 60 A) cause the energies and current densities of these jet ions to increase substantially. A model describing jet ion generation is proposed. The effects of discharge current on cathode internal pressure are also examined experimentally and described phenomenologically.

Navier-Stokes solution of transonic cascade flows using nonperiodic C-type grids

Abstract: A new kind of C-type grid is proposed, this grid is non-periodic on the wake and allows minimum skewness for cascades with high turning and large camber. Reynolds-averaged Navier-Stokes equations are solved on this type of grid using a finite volume discretization and a full multigrid method which uses Runge-Kutta stepping as the driving scheme. The Baldwin-Lomax eddy-viscosity model is used for turbulence closure. A detailed numerical study is proposed for a highly loaded transonic blade. A grid independence analysis is presented in terms of pressure distribution, exit flow angles, and loss coefficient. Comparison with experiments clearly demonstrates the capability of the proposed procedure.

Multiple Normal Shock Wave/Turbulent Boundary-Layer Interactions

Abstract: A multiple normal shock/turbulent boundary-layer interaction in a rectangular duct has been investigated experimentally using a two-component laser Doppler velocimeter (LDV). Just upstream of the normal shock system the Mach number was 1.61, the unit Reynolds number was 30 x 106 m ', and the confinement level as characterized by the ratio of the boundary-layer thickness to the duct half height was 0.32. The results presented here identify the fluid dynamic mechanisms involved in the reacceleration process following each shock in the multiple shock system. Two distinct expansion processes occur following each shock in the system: acceleration through a supersonic expansion fan originating near the wall and acceleration through an aerodynamic converging-diverging nozzle in the core flow.

State-feedback control of longitudinal combustion instabilities

Abstract: Active control of longitudinal pressure oscillations in combustion chambers has been studied theoretically by means of a digital state-feedback control technique. The formulation is based on a generalized wave equation that accommodates all influences of combustion, mean flow, unsteady motions, and control actions. Following a procedure equivalent to the Galerkin method, a system of ordinary differential equations governing the amplitude of each oscillatory mode is derived for the controller design. The control actions are provided by a finite number of point actuators, with instantaneous chamber conditions monitored by multiple sensors. Several important control aspects, including sampling period, locations of sensors and actuators, controllability, and observability, have been investigated systematically. As a specific example, the case involving two controlled and two residual (uncontrolled) modes is studied. 20 refs.

Rotational inviscid flow in laterally burning solid-propellant rocket motors

Abstract: A theoretical analysis to determine the effects of mass addition on the inviscid but rotational and compressible flowfield in a porous duct with the injection rate dependent on the local pressure is performed for large ratios of length-to-duc t diameter. The problem of describing the flow is reduced to the solution of a single integral equation. The ratio of specific heat 7, and a constant pressure exponent u, measuring the dependence of the rate of mass injection on the local pressure, are the parameters of the solutions. The integral equation is solved numerically, and parametric results are presented for 7, varying from 1 to f and for n varying from 0 to 1. A choking phenomenon is exhibited at a critical length of the duct in the vicinity of which the Mach number approaches unity. The choking condition, which is relevant to the operation of nozzleless solid-propellant rocket motors, is obtained parametrically in the present study and compared with corresponding results for irrotational, quasi-one-dimensional flow. The rotationality reduces the choking pressure. Nomenclature A = cross-sectional area of the duct a = radius of the duct C = average speed-of-sound in the fluid cp = the specific heat of the fluid at constant pressure h = half-height of channel 7sp = specific impulse of the motor K = constant prefactor in the burning-rate law for the propellant / = length of the duct M = Mach number m = mass burning rate of the propellant n = pressure exponent in the burning-rate law for the propellant P = nondimensional pressure p/pQ p = dimensional pressure R - ratio of the specific impulse obtained by assuming quasi-one-dimensional flow to the specific impulse obtained by assuming rotational flow r = dimensional radial coordinate 5 = perimeter of the duct T - temperature t = time u = axial component of the velocity Vw = injection velocity v = radial component of the velocity X — nondimensional axial strained coordinate defined in Eq. (23) x = dimensional axial coordinate y = distance normal to.the propellant surface y = ratio of specific heat A =. boundary-layer thickness 8 = gas-phase flame standoff distance 17 = dimensional transverse coordinate in channel flow p pp = gas density = propellant density = function related to entropy

Spectrally resolved Rayleigh scattering diagnostic for hydrogen-oxygen rocket plume studies

Abstract: A Rayleigh scattering diagnostic has been developed to measure gas density, temperature, and velocity in the exhaust plume of 100 N thrust class hydrogen-oxygen rockets. The spectrum of argon-ion laser light scattered by the gas molecules in the plume (predominantly water vapor) is measured with a scanning Fabry-Perot interferometer. The gas density is determined from the total scattered power, the gas temperature from the spectral width, and the velocity from the shift in the peak of the spectrum from the frequency of the incident laser light. The diagnostic has been demonstrated in a rocket test cell and a discussion of results is given.

Evaluation of parallel injector configurations for Mach 2 combustion

Abstract: An experimental study of wall-mounted parallel injector ramps has been conducted to explore techniques to enhance mixing in a scramjet combustor. Downstream parallel injection may be useful at high speeds to extract thrust from hydrogen that has been used to cool the engine and the airframe. The swept ramp fuel injector employed here should produce vortex shedding and local separation (like a rearward-facing step), which should enhance mixing. Perpendicular fuel injectors were added downstream of the swept ramps to determine if the vortical wake flow generated by the parallel injectors—with no fuel injection—is effective in enhancing the mixing of the transverse fuel jet. For performance comparisons, an unswept, but otherwise identical, parallel injector was also tested. The injector ramps were designed to yield a reflected shock wave from the duct top wall such that it passed just downstream of the barrel shock of the fuel injectors. The Mach number at the exit of the fuel injector was designed to be 1.7 to produce an underexpanded fuel flow under almost all operating conditions. Direct-connect tests were conducted with a vitiated heater over a total temperature range. Flow visualization tests with an unducted freejet configuration were conducted using shadowgraph and ultraviolet television camera systems for OH radical visualization. Three duct configurations were tested. Tests of the swept ramp with a 100-mm constant-area duct downstream of the injector block resulted in upstream interaction at almost any equivalence ratio. Increasing the duct expansion rate allowed the equivalence ratio to be increased beyond 1.5 with rapid mixing and combustion over a wide total temperature range. The unswept ramp design resulted in lower combustion efficiency and a sensitivity of efficiency to heater total temperature, or duct velocity.

Liquid fuel jet in subsonic crossflow

Abstract: An analytical/numerical model is described which predicts the behavior of nonreacting and reacting liquid jets injected transversely into subsonic cross flow. The compressible flowfield about the elliptical jet cross section is solved at various locations along the jet trajectory by analytical means for free-stream local Mach number perpendicular to jet cross section smaller than 0.3 and by numerical means for free-stream local Mach number perpendicular to jet cross section in the range 0.3-1.0. External and internal boundary layers along the jet cross section are solved by integral and numerical methods, and the mass losses due to boundary layer shedding, evaporation, and combustion are calculated and incorporated into the trajectory calculation. Comparison of predicted trajectories is made with limited experimental observations.

Electromagnetic effects in an applied-field magnetoplasmadynamic thruster

Abstract: Experimental and analytical studies have been conducted on the performance and thrust production mechanisms of an applied-field magnetoplasmadynamic thruster. The thruster was able to run with a high-thruster performance due to large electromagnetic effects related to the applied magnetic field. Using hydrogen, helium, and argon as the propellant, over 20 percent thrust efficiency was obtained over a wide specific impulse range from 1000 to 7000 s at input power levels between 2.2 and 15.9 kW. From the measurements of performance characteristics and current densities in the acceleration region, and by a theoretical analysis, it is found that the thruster operation is characterized by a parameter, B-squared/m (B: applied magnetic field strength, m: propellant mass flow rate). 9 refs.

Effectiveness of plasma torches for ignition and flameholding in scramjet

Abstract: A newly developed plasma torch with a feed stock of air or oxygen was investigated experimentally in order to determine its effectiveness on ignition and flameholding in a scramjet combustor. This design comes from the viewpoint of total system design of scramjet engine and vehicle because it is preferable to utilize incoming air or onboard propellants as a feed stock. Three patterns of fuel injection were tested 1) from one orifice: 2) from four orifices on one wall; and 3) from all nine orifices on both walls. Ignition and flameholding phenomena were examined through direct photographs of internal and exit flames of the combustor and by wall temperature measurements. The specially devised plasma torch with air or oxygen was able to operate stably without any support gas, for example, argon. Ignition limit curves, with and without the plasma torch, were obtained on a plane relating the air total temperature to the fuel equivalence ratio for the three patterns of fuel injection, and then compared to each other. For a wide range of experimental conditions, it was shown that the effectiveness of an air or oxygen plasma torch was comparable to that of a nitrogen or argon-hydrogen plasma torch. For single-wall injection, it was observed that the plasma torch ignited the fuel jet located directly downstream, and the flame thus formed ignited adjacent fuel jets. In double-wall injection, however, ignition of the fuel injected from the wall opposite the plasma torch was unsuccessful. It was also found that, under some conditions, flameholding can be continued even after the plasma torch is turned off, most notably in the case of single-wall injection. The occurrence or nonoccurrence of this phenomenon is also shown in the ignition limit curves diagram.

Wave combustors for trans-atmospheric vehicles

Abstract: A performance analysis is given of a conceptual transatmospheric vehicle (TAV). The TAV is powered by a an oblique detonation wave engine (ODWE). The ODWE is an airbreathing hypersonic propulsion system which utilizes shock and detonation waves to enhance fuel-air mixing and combustion in supersonic flow. In this wave combustor concept, an oblique shock wave in the combustor can act as a flameholder by increasing the pressure and temperature of the air-fuel mixture, thereby decreasing the ignition delay. If the oblique shock is sufficiently strong, then the combustion front and the shock wave can couple into a detonation wave. In this case, combustion occurs almost instantaneously in a thin zone behind the wave front. The result is a shorter lighter engine compared to the scramjet. The ODWE-powered hypersonic vehicle performance is compared to that of a scramjet-powered vehicle. Among the results outlined, it is found that the ODWE trades a better engine performance above Mach 15 for a lower performance below Mach 15. The overall higher performance of the ODWE results in a 51,000-lb weight savings and a higher payload weight fraction of approximately 12 percent.

Ram accelerator direct space launch system - New concepts

Abstract: The ram accelerator, a chemically driven ramjet-in-tube device is a new option for direct launch of acceleration-insensitive payloads into earth orbit. The projectile is the centerbody of a ramjet and travels through a tube filled with a premixed fuel-oxidizer mixture. The tube acts as the cowl of the ramjet. A number of new concepts for a ram accelerator space launch system are presented. The velocity and acceleration capabilities of a number of ram accelerator drive modes, including several new modes, are given. Passive (fin) stabilization during atmospheric transit is investigated and found to be promising. Gasdynamic heating in-tube and during atmospheric transit is studied; the former is found to be severe, but may be alleviated by the selection of the most suitable drive modes, transpiration cooling, or a hydrogen gas core in the launch tube. To place the payload in earth orbit, scenarios using one impulse and three impulses (with an aeropass) and a new scenario involving an auxiliary vehicle are studied. The auxiliary vehicle scenario is found to be competitive regarding payload, and requires a much simpler projectile, but has the disadvantage of requiring the auxiliary vehicle.

Metallized Propellants for the Human Exploration of Mars

Abstract: Advanced chemical propulsion using Metallized Propellants (MP) can lead to significant reductions in launch mass for piloted Mars missions. MP allow the propellant density or the specific impulse I(sub sp) of the propulsion system, or both, to increase. It can reduce the propellant mass and the propulsion system dry mass. Detailed mass-scaling equations and estimates of the I(sub sp) for several MP combinations are presented. The most significant savings with MP are derived from increasing the payload delivered to Mars. For the mass in low Earth orbit (LEO), a metallized Mars transfer vehicle can deliver 20 to 22 percent additional payload. This 20-percent payload increase reduces the total number of Mars flights and therefore significantly reduces the number of Space Transportation System-Cargo launches for the entire Mars architecture. Using MP to reduce the mass in LEO per flight is not as effective as increasing the payload delivery capacity. The mass saving per flight, while delivering the same payload with a higher I(sub sp) system, is much smaller. Using MP in all of the Mars propulsion systems would produce a modest 3.3 percent LEO mass saving. This translates into a saving of 38,000 kg over the mass required with O2/H2 propulsion. A Mars excursion vehicle using Earth- or space-storable propellants for the ascent can be an alternative to storing cryogenic H2 on Mars. A space-storable system using oxygen/monomethyl hydrazine/aluminum (O2/MMH/Al) would deliver the lowest mass penalty over O2/H2. For lower-energy expedition missions the LEO mass penalty for using metallized O2/MMH/Al would be only 3 to 5 percent.

Directions in Turbulence Modeling for In-Cylinder Flows in Reciprocating Engines

Abstract: Turbulence modeling is an essential aspect of computational fluid dynamics (CFD) for applications of interest in the automotive industry. In this paper, turbulence models are reviewed in the context of in-cylinder flows in reciprocating engines. While the ubiquitous k — e remains the model of choice in most applications, higherorder statistical closures and large-eddy simulation have been brought to bear in an effort to increase the predictive capability of multidimensional simulations for these complex flows. The performance of several models in simple engine-like configurations is reviewed and applications to flows of more practical interest are summarized. An assessment of the demonstrated and/or anticipated performance of different modeling approaches is given and suggestions of fruitful paths for further improvement are offered. The nature of turbulence in engines and numerical accuracy are two issues that are identified as being critical in multidimensional modeling of turbulent in-cylinder flows. Probability density function methods and large-eddy simulation are proposed as candidates for next-generation "turbulence models" that have the potential for significant improvements over today's k — £ based approaches.

Computational studies of a superdetonative ram accelerator mode

Abstract: A computational study of a ramjet-in-tube concept known as the "ram accelerator'9 is presented. The analysis assumes inviscid flow, and employs a total variation diminishing (TVD) numerical scheme that includes nonequilibrium chemistry, real gas effects, and a 7-species, 8-reaction combustion model for hydrogen/oxygen mixtures. This numerical formulation gives a more accurate model for the combustion process and is computationally more efficient than previous numerical studies conducted on the ram accelerator. The flow, combustion, and performance characteristics of the ram accelerator are investigated for several projectile configurations in the superdetonative velocity range of 5.0-10.0 km/s. The distribution of various physical quantities along the ram accelerator, as well as temperature contours, are presented. Plots of ballistic efficiency and thrust pressure ratio are also included. Ballistic efficiencies of up to 28% and thrust pressure ratios as high as 17% are obtained. The results also show that efficient acceleration of projectiles is possible through velocities as high as 9 km/s.

Combustion characteristics of AP/HTPB propellants with burning rate modifiers

Abstract: Fe2O3, copper chromite catalyst (CC), LiF, and CaCO3 are the four burning rate modifying additives considered in this study. A wide variety of subatmospheric flames are observed for the propellants with additives. Under subatmospheric pressures a dark zone is present; this zone becomes thicker at lower pressure and with finer ammonium perchlorate (AP) particles. A smoldering surface without a gas-phase flame is noted for the additive Fe2O3. From the smoke deposits, obtained from the subatmospheric burning of propellants with burning rate enhancers, it is shown that the breakdown of heavy fuel molecules is better with finer AP particles. The low-pressure deflagration limit is reduced by the addition of Fe2O3 or CC and increased by the addition of LiF or CaCO3. In the presence of LiF or CaCO3, above certain additive concentration levels, reduction in AP particle size is found to depress the burning rate further. LiF, below certain concentration levels, acts as a burning rate enhancer up to a certain pressure and as a depressant thereafter. The type of additive influences the pressure at which the shape of burning AP surface changes from convex to concave and also the pressure at which the AP surface recesses relative to the mean surface.

Experimental investigation of terminal shock sensors for mixed-compression inlets

Abstract: This paper describes experimental investigations of devices designed for the nonintrusive detection of terminal shock location in mixed-compression inlets at high supersonic flight speeds. Systems based on sensing wall pressures by an array of wall-mounted transducers were selected for detailed study. Pressure signals were processed by three different methods: (1) interpretation of instantaneous pressure distributions, (2) detection of the turbulent intensity amplification occurring at the shock, and (3) determination of the upstream limit to which a search-tone, introduced at the downstream end of the channel, can propagate. The first two of these methods were tested in real time. The third method appeared feasible for weak shocks only; at high shock strengths, propagation upstream of the source could not be detected.

Flow establishment in a generic scramjet combustor

Abstract: The establishment of a quasi-steady flow in a generic scramjet combustor was studied for the case of a time varying inflow to the combustor. Such transient flow is characteristic of the reflected shock tunnel and expansion tube test facilities. Several numerical simulations of hypervelocity flow through a straight duct combustor with either a side wall step fuel injector or a centrally located strut injector are presented. Comparisons were made between impulsively started but otherwise constant flow conditions (typical of the expansion tube or tailored operations of the reflected shock tunnel) and the relaxing flow produced by the 'undertailored' operations of the reflected shock tunnel. Generally the inviscid flow features, such as the shock pattern and pressure distribution, were unaffected by the time varying inlet conditions and approached steady state in approx. the times indicated by experimental correlations. However, viscous features, such as heat transfer and skin friction, were altered by the relaxing inlet flow conditions.

Medium-frequency impulsive-thrust-excited slosh waves during propellant reorientation with a geyser

Abstract: Slosh wave excitation induced by a resettling flowfield activated by 1.0-Hz impulsive thrust during the course of liquid reorientation with the initiation of geyser for liquid-fill levels of 30, 50, 65, 70, and 80 percent has been studied. Characteristics of slosh waves of various frequencies excited by the resettling flowfield are discussed. Slosh wave excitations shift the fluid mass distribution in the container which imposes time-dependent variations in spacecraft moment of inertia. This information is important for spacecraft control during the course of liquid reorientation.

Investigation of oblique shock/boundary-layer bleed interaction

Abstract: The detailed flowfield characteristics in an oblique shock-wave/laminar-boundary-layer interaction with bleed were investigated. The numerical solution for the flowfield was obtained for the strong conservation-law form of the two-dimensional compressible Navier-Stokes equations using an implicit scheme. The computations mod- eled the flow in the interaction region and inside the bleed slot for an impinging oblique shock on a flat-plate boundary layer. The computed results for the streamlines and the pressure and Mach number contours inside the bleed slot indicate that the flow is choked in the slot, with a recirculation zone near the upstream slot corner. The bleed results in the interaction zone demonstrate that flow separation is controlled. The interaction length . is reduced and the downstream velocity profiles are more favorable than the separated flow results at the same shock strength without bleed. HE control of shock/boundary-layer interactions in inlets and nozzles and over vehicle surfaces is accomplished through bleed and/or blowing in the interaction zone. In the case of mixed compression supersonic inlets, the bleed system design is critical to the efficient and stable operation of the system. Hamed and Shang1 reviewed the existing experimen- tal data for shock-wave/boundary-layer interactions in super- sonic inlets and other related configurations. According to this survey, most of the experimental measurements in mixed compression supersonic inlets consisted of total pressure re- covery surveys at the engine face and static pressure distri- butions over the inner surfaces. In the few cases involving velocity profile measurements,2-3 the latter were obtained up- stream and downstream of the interactions. Comparisons of internal flow computational results3'5 with the experimental measurements in supersonic inlets2-3 revealed reasonable agreement between the computed and measured surface pres- sures upstream of the ramp bleed. However, discrepancies in the predicted shock locations and velocity profiles were ob- served downstream of shock/boundary-layer interactions with bleed. There is enough experimental evidence6'11 to indicate that local bleed can control flow separation in shock-wave/bound- ary-layer interactions. There are disagreements,1 however, among the different experimental studies regarding the effects of bleed hole size,7-8 and the location of the bleed holes in relation to the shock.6-9'11 The experimental data in these studies are not sufficient, however, to resolve these discrep- ancies. Strike and Rippy9 measured the surface pressure in the interaction zone of an oblique shock wave impinging a tur- bulent boundary layer over a flat plate, with suction. They determined that less suction is required to control separation, when applied upstream of the shock. Seebaugh and Childs11 investigated experimentally the axisymmetric flow in the in- teraction region of the boundary layer inside a duct. Contrary to the conclusions of Strike and Rippy,9 suction within the

Ammonium Nitrate-Magnesium Propellant Combustion and Heat Transfer Mechanisms

Abstract: Microthermocouples and fiber optics were used to investigate the mechanisms of heat transfer and combustion in magnesium-ammonium nitrate (AN) composite propellants. It was found that a liquid layer at 300 ± 30°C covers the surface of both the AN and binder regions during combustion. The magnesium burns near the propellant surface in a dual vapor and surface mode producing both fine smoke and large-scale MgO ash. With no magnesium, conductive heat feedback from the partially premixed AN/binder gas flame drives the combustion of the propellant. The temperature of this flame varies with space and time, ranging between 1000°C and 160()°C. The spatial and temporal variations of the local flame temperature correspond to the local concentration fluctuations of the gas mixture above the propellant, which, in turn, correlate with the macroscopic heterogeneity established by the coarse AN prills and the binder-fine-AN-fill region. Burning rate was found to increase significantly with increasing magnesium loading due to conductive and radiative heat feedback from burning magnesium particles that were retained near the propellant surface by large-scale ash and due to condensedphase heat release from magnesium oxidation.

Solid-propellant erosive burning

Abstract: A modeling approach based on experimental verification for the erosive burning of ammonium perchlorate composite propellants is proposed. It retains the idea that erosive burning is due to the penetration of the flow turbulence within the flame height and to the consequent enhancement of the transport coefficients and of the heat flux to the surface. It aims at matching the levels of the models of the flow (a numerical shear-layer computation) and of the flame (a realistic diffusion flame applicable to large- and medium-size AP composite propellants—those most sensitive to erosive burning). Both the flow description and the computed erosive burning are successfully compared to experimental results: cold flow-porous wall simulation device with laser anemometry-measured velocity profiles and propellant erosive burning measured with an ultrasonic transducer on a separated generator-sample device. F h h

Application of an unsteady navier-stokes solver to transonic turbine design

Abstract: This study presents a numerical evaluation of the performance of the first stage of a new-generation turbine design. The numerical method solves the two-dimensional Navier-Stokes equations using a system of patched grids. Three-dimensional effects of stream-tube contraction are also modeled. The study focuses on the effects of axial gap variation on the unsteady rotor-stator interactions and on stage performance. Results are presented for three different axial gaps. The results indicate that the unsteady interactions can be very large in this design. These interactions affect not only the stage efficiency but also substantially alter the time-averaged features of the flow. In particular, for the case of the smallest axial gap, it was found that there was an unsteady shock on the stator suction surface which spanned the gap region and impinged upon the moving rotor airfoils.