Showing papers in "Journal of Propulsion and Power in 2010"

Review: Laser-Ablation Propulsion

Abstract: LASER ablation propulsion (LAP) is a major new electric propulsion concept with a 35-year history. In LAP, an intense laser beam [pulsed or continuous wave (CW)] strikes a condensedmatter surface (solid or liquid) and produces a jet of vapor or plasma. Just as in a chemical rocket, thrust is produced by the resulting reaction force on the surface. Spacecraft and other objects can be propelled in this way. In some circumstances, there are advantages for this technique compared with other chemical and electric propulsion schemes. It is difficult to make a performance metric for LAP, because only a few of its applications are beyond the research phase and because it can be applied in widely different circumstances that would require entirely different metrics. These applications range from milliwatt-average-power satellite attitude-correction thrusters through kilowatt-average-power systems for reentering near-Earth space debris and megawatt-to-gigawatt systems for direct launch to lowEarth orbit (LEO). We assume an electric laser rather than a gas-dynamic or chemical laser driving the ablation, to emphasize the performance as an electric thruster. How is it possible for moderate laser electrical efficiency to givevery high electrical efficiency? Because laser energy can be used to drive an exothermic reaction in the target material controlled by the laser input, and electrical efficiency only measures the ratio of exhaust power to electrical power. This distinction may seem artificial, but electrical efficiency is a key parameter for space applications, in which electrical power is at a premium. The laser system involved in LAP may be remote from the propelled object (on another spacecraft or planet-based), for example, in laser-induced space-debris reentry or payload launch to low planetary orbit. In other applications (e.g., the laser–plasma microthruster that we will describe), a lightweight laser is part of the propulsion engine onboard the spacecraft.

Simultaneous Pressure and Optical Measurements of Nanoaluminum Thermites: Investigating the Reaction Mechanism

Abstract: This work investigates the reaction mechanism of metastable intermolecular composites by collecting simultaneous pressure and optical signals during combustion in a constant-volume pressure cell. Nanoaluminum and three different oxidizers are studied: CuO, SnO2, and Fe2O3. In addition, these mixtures are blended with varying amounts of WO3 as a means to perturb the gas release in the system. The mixtures with CuO and SnO2 exhibit pressure signals that peak on timescales faster than the optical signal, whereas themixtures containingFe2O3 do not show this behavior. The burn time is found to be relatively constant for bothCuOandSnO2, evenwhen a large amount ofWO3 is added. For Fe2O3, the burn time decreases asWO3 is added, and the temperature increases. The results are consistent with the idea that oxidizers such as CuO and SnO2 decompose and release gaseous oxidizers fast, relative to the burning, and this is experimentally seen by an initial pressure rise followed by a prolonged optical emission. In this case, the burning is rate limited by the aluminum, and it is speculated to be similar to the burning of aluminum in a pressurized oxygenated environment. For the Fe2O3 system, the pressure and optical signals occur concurrently, indicating that the oxidizer decomposition is the rate-limiting step.

Experimental Study of Cavity-Strut Combustion in Supersonic Flow

Abstract: An experimental investigation of cavity-based flameholders with strut injectors in a supersonic flow is reported. In this ongoing research program, emphasis is placed on understanding cavity-based flameholders and providing alternative methods for improving overall combustor performance in scramjet engines. Three different struts with fuel injectors are mounted near the cavity leading edge to study flame propagation and ignition of fuel in the core flow region. Planar laser-induced fluorescence of the OH radical is used to identify the flame zone around the cavity and strut-wake regions over a range of conditions. Shadowgraphy is employed to capture the flow features around the strut and cavity. In-stream probing is conducted to characterize the flow features associated with the different strut configurations. Stagnation-temperature profiles are obtained for all struts operating under the same condition in the combusting-flow study. Two cavity fueling schemes are used to compare flameholder performance. Direct cavity air injection is found to improve combustion significantly. For each strut, upstream and downstream fueling schemes are compared over a range of conditions. Overall, successful combustion is observed in the strut-wake region using upstream strut-fueling schemes for the three struts employed in this study.

Numerical investigation of transverse jets through multiport injector arrays in supersonic crossflow

Abstract: A three-dimensional numerical study has been performed of the effects of sonic gaseous hydrogen injection through multiple transverse injectors subjected to a supersonic crossflow. Solutions were obtained for a series of injection configurations in a Mach 4.0 crossflow, with a global equivalence ratio of o = 0:5. Results indicate a different flow structure than for a typical single jet, with the development of two clearly defined wake vortices, including a stagnation point and reversed flow region immediately behind each downstream jet. While the overall penetration was reduced under the investigated conditions, significant improvements were observed when nondimensionalizing against the equivalent jet diameter for each modeled injector row. This was found to be the result of increased jet-to-freestream momentum ratio due to the subsonic flow regions between each injector. Further enhancements were also observed in terms of mixing performance for the multijet cases. Improvements of up to 5% in the overall mixing efficiency were experienced by using multiple jets due to increased mixant interface area and intermediate stirring through wake vortices between each injector. No improvement in far-field mixing was observed. Overall, it has been demonstrated that there are benefits to be gained through the injection of gaseous hydrogen from many small injectors rather than fewer large injectors. Copyright © 2010 by A. S. Pudsey and R. R. Boyce. Published by the American Institute of Aeronautics and Astronautics, Inc.

Acoustic Characterization of an Ethylene-Fueled Scramjet Combustor with a Cavity Flameholder

Abstract: *† ‡ § ** †† ‡‡ The occurrence of combustion oscillations has recently raised serious concerns about the development of scramjet engines. Previous studies on supersonic combustion for high-speed airbreathing propulsion applications indicated that combustion may take place in subsonic regions, such as boundary layers and recirculation zones in flame-holding cavities. During this process, a longitudinal mode of thermoacoustic instability may develop in a spatial domain reaching from the shock train to the flame zone. The present work experimentally and analytically investigates such thermoacoustic instabilities inside an ethylene-fueled scramjet combustor with a recessed cavity flameholder. High-speed pressure transducers are utilized to record acoustic signals. The effects of fuel/air equivalence ratio, fueling scheme, and simulated flight conditions on the stability characteristics of the combustor are examined systematically. A companion analytical analysis is also established to help explore the underlying mechanisms responsible for driving and sustaining thermoacoustic flow instabilities. In particular, the interactions between the unsteady heat release, fuel injection and mixing, and shock response are examined. The measured oscillation frequencies agree well with the characteristic frequencies related to the acoustic feedback loop between the shock and flame and the acoustic-convective feedback loop between the fuel injection and flame.

Energy Content and Alternative Jet Fuel Viability

Abstract: This paper examines the chemical composition and energy content of several fuel options that could hypothetically be used with the existing fleet of aircraft. Fuel specific energy (energy per unit mass) is an important consideration in determining alternative-fuel viability, because aircraft must travel fixed distances before refueling. Since most aircraft fly with excess tank capacity, fuel energy density (energy per unit volume) is of secondary concern relative to specific energy. A first-order approach using the Breguet-range equation shows that the fleet-wide use of pure synthetic paraffinic kerosene fuels, such as those created from Fischer-Tropsch synthesis or hydroprocessing of renewable oil sources, could reduce aircraft energy consumption by 0.3%. Conversely, fuels with reduced specific energy, such as fatty acid methyl esters (biodiesel and biokerosene) and alcohols, will result in increased fuel volume usage and also a decrease in fleet-wide energy efficiency. No penalty in energy efficiency would occur were these fuels used in ground transportation; thus, fatty acid methyl esters and alcohols are better suited to use in ground-based applications.

Liquid-propellant rocket engine throttling: A comprehensive review

Abstract: Liquid-propellant rocket engines are capable of on-command variable thrust or thrust modulation, an operability advantage that has been studied intermittently since the late 1930s. Throttleable liquid-propellant rocket engines can be used for planetary entry and descent, space rendezvous, orbital maneuvering including orientation and stabilization in space, and hovering and hazard avoidance during planetary landing. Other applications have included control of aircraft rocket engines, limiting of vehicle acceleration or velocity using retrograde rockets, and ballistic missile defense trajectory control. Throttleable liquid-propellant rocket engines can also continuously follow the most economical thrust curve in a given situation, as opposed to making discrete throttling changes over a few select operating points. The effects of variable thrust on the mechanics and dynamics of an liquid-propellant rocket engine as well as difficulties and issues surrounding the throttling process are important aspects of throttling behavior. This review provides a detailed survey of liquid-propellant rocket engine throttling centered around engines from the United States. Several liquid-propellant rocket engine throttling methods are discussed, including high-pressure-drop systems, dual-injector manifolds, gas injection, multiple chambers, pulse modulation, throat throttling, movable injector components, and hydrodynamically dissipative injectors. Concerns and issues surrounding each method are examined, and the advantages and shortcomings compared.

Ignition Delay Time Measurements of Normal Alkanes and Simple Oxygenates

Abstract: Ignition delay time data for higher n-alkanes are required for the development and refinement of jet fuel surrogate mechanisms; similar data for oxygen-carrying species are also required for the development of oxygenate fuel mechanisms. To fill this need, ignition delay times were measured using pressure and OH * emission diagnostics behind reflected shock waves for a series of higher n-alkanes and for several simple oxygenates. Reflected shock conditions covered temperatures of 1150―1550 K and pressures of 1―4 atm (0.1―0.4 MPa). Fuel mixtures tested include the four n-alkanes: n-pentane, n-hexane, n-octane, and n-nonane; and the four oxygenates: acetone, n-butanal, methyl butanoate, and 3-pentanone. All fuels were tested in oxygen/argon mixtures with equivalence ratios of 0.5 to 2.0. The ignition data were compared with current models and a new n-butanal reaction mechanism, based on previous studies, was used to model the n-butanal and methyl butanoate data.

Microstructure Effects in Aluminized Solid Rocket Propellants

Abstract: Experiments concerning the ballistic characterization of several nanoaluminum (nAl) powders are reported. Most studies were performed with laboratory composite solid rocket propellants based on ammonium perchlorate as oxidizer and hydroxyl-terminated polybutadiene as inert binder. The ultimate objective is to understand the flame structure of differently metallized formulations and improve their specific impulse efficiency by mitigating the two-phase losses. Ballistic results confirm, for increasing nAl mass fraction or decreasing nAl size, higher steady burning rates with essentially the same pressure sensitivity and reduced average size of condensed combustion products. However, aggregation and agglomeration phenomena near the burning surface appear noticeably different for microaluminum (μAl) and nAl powders. By contrasting the associated flame structures, a particle-laden flame zone with a sensibly reduced particle size is disclosed in the case of nAl. Propellant microstructure is considered the main controlling factor. A way to predict the incipient agglomerate size for μAl propellants is proposed and verified by testing several additional ammonium perchlorate/hydroxyl-terminated polybutadiene/aluminum formulations of industrial manufacture.

Reduced-Order Modeling of Two-Dimensional Supersonic Flows with Applications to Scramjet Inlets

Abstract: Control-oriented models of hypersonic vehicle propulsion systems require a reduced-order model of the scramjet inlet that is accurate to within 10% but requires less than a few seconds of computational time. To achieve this goal, a reduced-order model is presented, which predicts the solution of a steady two-dimensional supersonic flow through an inlet or around any other two-dimensional geometry. The model assumes that the flow is supersonic everywhere except in boundary layers and the regions near blunted leading edges. Expansion fans are modeled as a sequence of discrete waves instead of a continuous pressure change. Of critical importance to the model is the ability to predict the results of multiple wave interactions rapidly. The rounded detached shock near a blunt leading edge is discretized and replaced with three linear shocks. Boundary layers are approximated by displacing the flow by an empirical estimate of the displacement thickness. A scramjet inlet is considered as an example application. The predicted results are compared to two-dimensional CFD solutions and experimental results. Nomenclature a = local soundspeed [m/s] c = specific heat [J/kg K] h = specific enthalpy [J/kg] H = length normal to flow [m] M = Mach number n = number of a given quantity L = length tangent to flow [m] p = pressure [Pa] Pr = Prandtl number r = radius [m] R = normalized gas constant [J/kg K] R = 8314.47 J/kmol K T = temperature [K] u = velocity magnitude [m/s] W = molecular weight [kg/kmol] x = forward body-frame coordinate [m] Y = mass fraction z = vertical body-frame coordinate [m] = shock angle = ratio of specific heats = thickness of layer [m] " = ratio = flowpath angle = dynamic viscosity [kg/m s] = ln p0=p = += 2

Unstart Detection in a Simplified-Geometry Hypersonic Inlet-Isolator Flow

Abstract: Unstart detection techniques based on high-frequency pressure measurements made in a hypersonic inlet―isolator model are investigated. In this study, data that were acquired in a previous study of backpressure-induced unstart were examined. The data were acquired in a simplified-geometry inlet―isolator model that consisted of a 6 deg compression ramp inlet followed by a constant-area duct that was 25.4 mm high by 50.8 mm wide by 227.1 mm long. Fluctuating wall pressures were measured along the length of the model. A downstream flap was used to induce unstart in the model. Beyond a certain flap angle, unstart was induced, and the shock system propagated upstream and out of the inlet. The wall-pressure data, acquired as the flap was raised, were postprocessed for spectral and statistical content to evaluate different unstart detection criteria. Three shock leading-edge detection criteria are examined based on the following observations as the shock system passes over a pressure transducer: 1) a rise in pressure, 2) an increase in standard-deviation of the pressure signal, and 3) an increase in power in the 300―400 Hz frequency band. After calibrating the algorithms based on runs with no active control, a comparison of the times detected for unstart onset and unstart arrest was made based on runs with active control. Results indicate that the power-spectrum-based algorithm implemented close to the isolator exit is more sensitive to the onset of unstart, whereas the pressure-magnitude-change criterion gives earlier detection in many cases. A combination of the two, a pressure-magnitude criterion close to the inlet entrance and a spectral-power criterion near the isolator exit, therefore seems to be the most robust scheme for unstart active control. The appropriate choice of sampling frequency significantly improved the computation speed of the spectral-power-based algorithm without delaying the unstart detection times.

Electrostatic Plasma Brake for Deorbiting a Satellite

Abstract: Space debris in the form of abandoned satellites is a growing concern, especially at the heavily populated 600― 1000 km altitude orbits. To prevent new space junk from forming, new satellites should be equipped with a deorbiting mechanism. The problem is especially tricky for the emerging class of very small satellites for which using a braking rocket as a deorbit mechanism may have a prohibitively high relative cost impact. We describe a novel type of deorbiting mechanism that is suitable for small satellites with a mass of up to a few hundred kilograms. The method is a plasma brake device based on coulomb drag interaction between the ionospheric plasma and a negatively charged thin tether. The method resembles the well-known electrodynamic tether deorbit mechanism, but the underlying physical mechanism is different and the new method has an order of magnitude smaller mass and power consumption. The new method uses the same physical principle (coulomb drag) as the recently invented electric solar wind sail propulsion method. Furthermore, the tether required by the plasma brake is so thin that, if accidentally cut, the loose fragments of it pose no threat to other spacecraft and will rapidly descend into the atmosphere. The electrostatic plasma brake could enable an extended use of small satellites by resolving their associated space debris problem.

Experimental Study of Fluid Jet Mixing at Supercritical Conditions

Abstract: Subcritical and supercritical fluids were injected in an inert gaseous atmosphere. Density distribution was measured and density-gradient profiles were inferred from the experimental data. A novel method was applied for the detection of detailed structures throughout the entire jet center plane. The core lengths were measured for each of the cases and correlated with previous visualization results. An eigenvalue approach was taken to determine the location of maximum gradients. The results show a significant influence of chamber-to-injectant density ratio on the core length in the supercritical domain, unlike the subcritical conditions.

Raman Scattering Measurements of Gaseous Ethylene Jets in a Mach 2 Supersonic Crossflow

Abstract: The structures of sonic ethylene jets delivered from orifices of three different diameters and two injection angles (30 and 90deg) into aMach 2 supersonic crossflowwere studied experimentally. The ratio of the cross-sectional areas of the largest and smallest injectors is 25:1. Time-averaged spontaneous vibrational Raman scattering was used to quantify injectant concentrations by constructing two-dimensional spanwise concentration images from the onedimensional linewise Raman scattering images. Based on the present data set, new penetration height correlations were developed to treat cases with injection angles of both 30 and 90 deg. Excluding the influence of wall boundary layer, the present measurements show that the properties of fuel plume structures, such as shape, size, and concentration profiles, are scalable with the injector size. Themeasured ethylene concentrations were also compared with predictions from the revised jet penetration code, which was calibrated primarily with hydrogen and helium. Discrepancieswere observedbetween themeasurements and the jet penetration code predictions for the structures of ethylene fuel plumes. The experimental data generated from the present study can be used to validate the numerical simulations.

Combustion of Silicon/Teflon/Viton and Aluminum/Teflon/Viton Energetic Composites

Abstract: DOI: 10.2514/1.46182 The combustion of Si- and Al-based systems using polytetrafluoroethylene (PTFE) as the oxidizer and Fluorel FC 2175 (a copolymer of hexafluoropropylene and vinylidene fluoride) as a binder has been studied. Experimental data were obtained using two methods: 1) instrumented tube burns and 2) pressed pellets inside a windowed pressure vessel. Loose-powder burning rates were seen to optimize at slightly-fuel-rich mixture ratios for Si/PTFE/FC-2175 (SiTV). Al/PTFE/FC-2175 (AlTV) burning rates optimized near a stoichiometric ratio. Pressures calculated by assuming constant-volume combustion equilibrium were seen to match experimental values from burn-tube experiments when burning rates were at or near peak values. The pressure dependence of SiTV and AlTV pellet burning rates was also characterized and compared with reported Mg/PTFE/Viton (MTV) results. SiTV showed power-law dependence with a constant-pressure exponent over the experimental range of pressures. AlTV was showntoexhibitnonconstant-pressureexponentbehavior.SiTVburningratesoptimizedatmixtureratiossimilarto that of the tube burns. AlTV burning rates increased well past a stoichiometric ratio and decreased at a fuel-rich ratio, which is a similar trend to MTV burning rates.

Inlet Starting of High-Contraction Axisymmetric Scramjets

Abstract: Reliable in-flight starting of the inlet is of critical importance for the successful operation of scramjet engines, particularly axisymmetric configurations with high-contraction inlets. The present research is undertaken to examine the capability of various inlet starting methods based on two principles: unsteady flow effects and variable geometries. Time-accurate viscous computations have been performed to investigate the transitional flowfields introduced by a variety of methods that are applicable to axisymmetric geometries. Parametric studies have been conducted for instantaneous rupture of conical diaphragms and addition of bleed slots, which induce highly unsteady flow phenomena. Several methods employing variable inlet geometries have been tested for the latter principle, including opening doors and sliding doors (or diaphragm erosion). Successful inlet starting has been achieved as a result of unsteady transition induced by diaphragm rupture and quasi-steady transition, due to the sliding-door opening process. In particular, a bleed addition to the diaphragm rupture method has been found to be highly effective and pronounced flow stability has been observed in the sliding-door process.

Large-Eddy Simulation of Flame-Turbulence Interactions in a Shear Coaxial Injector

Abstract: Various state-of-the-art Computational Fluid Dynamics (CFD) approaches are currently being evaluated by NASA for the modeling of combustion chambers in liquid rocket engines. This evaluation is performed through the simulation of a laboratory-scale chamber equipped with a single-element gaseous hydrogen-gaseous oxygen GH2-GO2 shear-coaxial injector. While a joint paper 1 compares the dierent steady and unsteady simulations of this chamber, this paper focuses on the three-dimensional Large Eddy Simulations (LES) performed at the Georgia Institute of Technology. The goal here is to nd a compromise between accuracy and computational cost while accurately capturing the unsteady ame physics thanks to the full 3D formulation. The solver employed in this study includes a hybrid central-upwind scheme used to solve the compressible, multi-species Navier-Stokes equations with a thermally perfect formulation. From a resolution point of view, priority was given to the near-eld resolution rather than the near-wall resolution. Yet, the seemingly correct modeling of the ame anchoring and dynamics provides a satisfactory estimate of the wall heat ux. The unsteady and three dimensional features of the ow are discussed in length and the implications of the unique features of this shear-coaxial GH2-GO2 injector for turbulent combustion modeling are also analyzed.

Comparative Study of Spray Characteristics of Gas-Centered and Liquid-Centered Swirl Coaxial Injectors

Abstract: This study investigated the spray and Sauter mean diameter characteristics of gas-liquid swirl coaxial injectors by measuring and analyzing spray angles and mean drop sizes. Two different types of gas-liquid swirl coaxial injectors were designed and tested for this experiment: 1) a gas-centered swirl coaxial injector and 2) a liquid-centered swirl coaxial injector. The spray patterns were obtained for both types of injectors. The spray angles of these liquidcentered swirl coaxial injectors showed a decreasing trend as a function of the momentum flux ratio over the conditions studied. Good agreement was generally achieved between the predicted and measured spray angles. On the other hand, with increasing momentum flux ratio, the spray angle of the gas-centered swirl coaxial injector initially decreased when the momentum flux ratios were relatively low, but the trend reversed over the high momentum flux ratio range studied. The Sauter mean diameter data of these two types of injectors were also obtained using an image-processing method. Results showed that the Sauter mean diameter distribution of the liquid-centered swirl coaxial injector had a solid cone shape, whereas the gas-centered swirl coaxial injector exhibited two distinct Sauter mean diameter distribution regimes. The gas-centered swirl coaxial injector had a solid cone shape along the centerline, and it had a hollow cone shape at the periphery.

Experimental Study on Flow Transition in Dual Bell Nozzles

Abstract: The dual bell nozzle is a concept of altitude adaptive nozzles. The flow adapts to the altitude by separation at the wall inflection at low altitude, and full flowing at high altitude. To understand the phenomenology of the flow by the transition from sea level to high altitude mode, a series of tests have been made at the cold flow test bench P6.2 at DLR Lampoldshausen. Three nozzles with different geometries have been tested. Two of them were successively shortened and driven unter the same conditions for each extension length. This study yields the influence of the geometric parameters of the basis and the extension on the transition conditions. Furthermore a transition prediction is given.

Molecular Weight and Shock-Wave Effects on Transverse Injection in Supersonic Flow

Abstract: An experimental study of the effects of injectant molecular weight on transverse-jet mixing in a supersonic crossflow is reported. In addition, the effects of an oblique shock impinging near the injection station were investigated. The examined gaseous injector is circular in geometry and angled downstream at 30 deg to the horizontal. Test conditions involved sonic injection of helium, methane, and air at a jet-to-freestream momentum flux ratio of 2.1 into a nominal Mach 4 air cross stream with average Reynolds number 5.77e + 7 per meter to provide a range of injectant molecular weights from 4―29. Sampling probe measurements were used to determine the local helium and methane concentration. A miniature five-hole pressure probe, pitot and cone-static pressure probes, and a diffuser-thermocouple probe were employed to document the flow. The goals of this effort are twofold. The first goal is to broaden and enrich the database for transverse injection in high-speed flows. Second, these data will aid greatly in the development of advanced turbulence models with a wide range of applicability for high-speed mixing flows. The main experimental results showed that an impinging shock reduces penetration and increases mixing for injectants of all molecular weights. Higher molecular weight injectant seems to increase penetration, but the effect is weak. The effects of shock impingement were strongest when this occurred downstream of, and closest to, the injection port.

Simulation of Shock/Boundary-Layer Interactions with Bleed Using Immersed-Boundary Methods

Abstract: This work utilizes an immersed boundary (IB) method to simulate the effects of arrays of discrete bleed ports in controlling shock wave / turbulent boundary layer inter actions . Both Reynolds averaged Navier -Stokes (RANS) and hybrid large -eddy / Reynolds -averaged Navier -Stokes (LES/RANS) turbulence closures are used with the IB technique. The approach is validated by conducting simulations of Mach 2.5 flow over a perfo rated plate containing 18 individual bleed holes. Predictions of discharge coefficient as a function of bleed plenum pressure are compared with experimental data. Simulations of an impinging oblique shock / boundary layer interaction at Mach 2.45 with an d without active bleed control are also performed. The 68 -hole bleed plate is rendered as an immersed object in the computational domain. Wall pressure predictions show that, in general, the LES/RANS technique under -estimate s the upstream extent of axi al separation that occurs in the absence of bleed. Good agreement with P itot -pressure surveys throughout the interaction region is obtained, however. Active suction completely removes the separation region and induces local disturbances in the wall pres sure distributions that are associated with the expansion of the boundary layer fluid into the bleed port and its subsequ ent re -compression. Predicted Pitot -pressure distributions are in good agreement with experiment for the case with bleed. Swirl stre ngth probability -density distributions are used to estimate the evolution of turbulence length -sca les throughout the interaction, and the effects of bleed on the amplification of Reynolds stresses are highlighted. Finally, simple improvements to engineerin g-level bleed models are proposed based on the computational results.

Role of Tip Clearance Flow in Rotating Instabilities and Nonsynchronous Vibrations

Abstract: Single- and multiple-blade-passage simulations of an isolated subsonic axial compressor rotor show that flow oscillations in the tip region, known as rotating instabilities and a driver for nonsynchronous vibrations, occur when only one of the two criteria for short-length-scale rotating stall inception is satisfied. This criterion is tip clearance backflow below the trailing-edge blade tip. The flow oscillations associated with rotating instabilities most likely result from impingement of this tip clearance backflow on the rear pressure side of the blade. This phenomenon could plausibly be modeled with an impinging jet subject to a lateral pressure gradient and lateral shear flow. The findings have important practical implications on the prediction and suppression of nonsynchronous vibrations.

Particle Based Plasma Simulation for an Ion Engine Discharge Chamber

Abstract: A particle-based model with a Monte Carlo collision model has been developed to study the plasma inside the discharge chamber of an ion engine. This model tracks five major particle types inside the discharge chamber in detail: xenon neutrals, singly charged xenon ions, doubly charged xenon ions, secondary electrons, and primary electrons. Both electric and magnetic field effects are included in the calculation of the charged particle's motion. The electric fields inside the discharge chamber are computed using a new approach. Also, detailed particle collision mechanisms are enabled. Validation of this computational model has been made on NASA's three-ring Solar Electric Propulsion Technology Application Readiness Program discharge chamber, at the 2.29 kW input power, 1.76 A beam current, and 1100 V beam voltage operating condition. Comparisons of numerical simulation results with experimental measurements are found to have good agreement. The computed ion beam current differs from experiments by 1% and the computed discharge current differs from experiments by 22%. The plasma ion production cost compares within 7% and the beam ion production cost compares within 16% of the experimental values. The overall computed thruster efficiency is found to differ from experiments by 11 %. In addition, steady-state results are given for particle number density distributions, kinetic energy, particle energy loss mechanisms, and current density collected at the chamber walls.

Miniature Microwave Discharge Ion Thruster Driven by 1 Watt Microwave Power

Abstract: T HE age of space exploration using self-propelled small spacecraft is approaching. In the past decade, research and development of small spacecraft have advanced extensively throughout the world. Numerous small spacecraft have been successfully launched and operated [1–8]. Small-spacecraft missions are increasingly used with widely varying applications. In recent years, planned small-spacecraft missions are increasingly in need of propulsive capability. Propulsion devices supply the spacecraft with attitude control, station-keeping, and orbit transfer functions. Furthermore, propulsion can support future spacecraft missions such as drag-free control from atmospheric or solar pressure, precise constellation flight for interferometer missions, and deorbiting maneuvers of end-of-life spacecraft into the atmosphere. The arrival of propulsion devices suitable for small spacecraft, namely, micropropulsion, is eagerly anticipated [9]. Ion thrusters [10] are promising propulsion devices intended for use not only with standard-sized spacecraft but also with small spacecraft. Their characteristics of high specific impulse (3000 s), high thrust efficiency (50%), usage of an inert propellant (xenon), and continuously controllable thrustmeet the requirements for smallspacecraft missions. Several studies of miniature ion thrusters have been reported for plasma generators of several different types, namely, direct-current electron discharge [11,12], radio frequency discharge [13], and microwave discharge [14–17]. Among these types, microwave discharge ion thrusters have the following unique characteristics suitable for miniaturization. First, the discharge chamber generating the plasma has a simple structure because of its lack of an electron cathode. In this type, electrons are heated using electron cyclotron resonance (ECR); neither a discharge cathode generating primary electrons nor corresponding power supplies are needed. Second, the thruster performance is degraded only slightly by scaling down. In general, discharge loss is affected by the surfaceto-volume ratio of plasma. In microwave discharge ion thrusters, however, the ECR heating zone is confined to the near-surface magnetic field, meaning that the surface-to-volume ratio is not sensitive to the system’s scale. Using these advantages, Takao et al. [14], Yamamoto et al. [15,17], and Nakayama et al. [16] developed miniature ion thrusters driven by microwave discharges and showed good performance. In spite of these benefits, however,miniature ion thrusters have not been used for small spacecraft yet. An important reason is the severe limitation of electrical power available on small spacecraft. In general, the electrical power generated in small spacecraft is about 1 W=kg. For example, the maximum usable power in a 10–50 kg small spacecraft was estimated at only 10–50 W. The miniature ion thrusters developed to date require a total power of at least 30W.Low power limitations have made it difficult to use those ion thrusters on small spacecraft. On the other hand, mass limitations are not so severe problem as that of the electrical power. Nakayama et al. [16] estimated the total dry weight as approximately 4 kg for their miniature microwave discharge ion thruster. Additionally, this weight can be reduced father by decreasing themicrowave power and discharge loss because the microwave power supply has a large fraction of its total weight attributable to the low energy conversion efficiency. Our objective is a miniature ion thruster with a total power consumption of less than 10 W and a total dry weight of less than 2.0 kg. One such miniature ion thruster can propel a 10–20 kg small spacecraft, although several thrusters could propel larger spacecraft of 20–100 kg. Moreover, devices that produce minute and controllable thrust are useful for spacecraft that require precise control of both attitude and position. The major challenge to achieving this goal is to decrease the total required electrical power while maintaining high thruster performance. That is to say that power for generating plasma should be limited to only 1–2Wand the discharge loss should be suppressed to less than 500 W=A. To date, no ion thruster has been reported to operate at such lowpower applied to the plasma and at such a low discharge loss. As a first step toward this goal, we have developed a miniature ion thruster driven by 1 W of microwave power with a 250 W=A ion production cost and 37% mass utilization efficiency. In this paper, we present the characteristics and performance of aminiaturemicrowave dischargebased ion thruster.

Scalar Filtered Density Function for Large Eddy Simulation of a Bunsen Burner

Abstract: The scalar filtered density function methodology is employed for large eddy simulation of a turbulent stoichiometric premixed methane-air flame. The scalar filtered density function accounts for the subgrid-scale chemical reaction by considering the mass-weighted probability density function of the subgrid-scale scalar quantities. A transport equation is derived for the scalar filtered density function in which the effects of chemical reactions appear in closed form. The subgrid-scale mixing is modeled via the linear mean square estimation model, and the convective fluxes are modeled via a subgrid-scale viscosity. The modeled scalar filtered density function transport equation is solved by a hybrid finite difference and Monte Carlo scheme. A novel irregular Monte Carlo portioning procedure is developed that facilitates efficient simulations with realistic flow parameters. Combustion chemistry is modeled via five-step, nine-species reduced chemical kinetics. Simulated results are assessed by comparisons against laboratory data. Good agreements are observed, capturing several important features of the flame as observed experimentally.

Emissions Characteristics of Military Helicopter Engines with JP-8 and Fischer-Tropsch Fuels

Abstract: The rapid growth in aviation activities and more stringent U.S. Environmental Protection Agency regulations have increased concerns regarding aircraft emissions, due to their harmful health and environmental impacts, especially in the vicinity of airports and military bases. In this study, the gaseous and particulate-matter emissions of two General Electric T701C engines and one T700 engine were evaluated. The T700 series engines power the U.S. Army's Black Hawk and Apache helicopters. The engines were fueled with standard military JP-8 fuel and were tested at three power settings. In addition, one of the T701C engines was operated on a natural-gas-derived Fischer-Tropsch synthetic paraffinic kerosene jet fuel. Test results show that the T701C engine emits significantly lower particulate-matter emissions than the T700 for all conditions tested. Particulate-matter mass emission indices ranged from 0.2-1.4 g/kg fuel for the T700 and 0.2-0.6 g/kg fuel for the T701C. Slightly higher NOx and lower CO emissions were observed for the T701C compared with the T700. Operation of the T701C with the Fischer-Tropsch fuel rendered dramatic reductions in soot emissions relative to operation on JP-8, due primarily to the lack of aromatic compounds in the alternative fuel. The Fischer-Tropsch fuel also produced smaller particles and slight reductions inmore » CO emissions.« less

Numerical investigation of the supersonic combustion of kerosene in a strut-based combustor

Abstract: In this numerical study, supersonic combustion of liquid kerosene in a strut-based combustor is investigated. To this end, three-dimensional compressible, turbulent, nonreacting and reacting flow calculations with a single-step chemistry model have been carried out. For the nonreacting flow calculations, fuel droplet trajectories, degree of mixing,andmixingefficiencyarepresentedanddiscussed.Forthereacting flowcalculations,contoursofheatrelease and Mach number and the variation of combustion efficiency, total pressure loss, and thrust profile along the combustor length are used to identify the regions of mixing and heat release inside the combustor. Furthermore, the predicted variation of static pressure along the combustor top wall is compared with experimental data. The significance of the lateral spread of the fuel and the extent of the mixing process, especially for a liquid fuel such as kerosene, on the prediction of heat release is discussed in detail.

Linear Response of Laminar Premixed Flames to Flow Oscillations: Unsteady Stretch Effects

Abstract: �c and � � s ,forcurvatureand hydrodynamic strain, respectively Unsteady curvature effects on the flame surface area become significant when j� � c jSt 2 2 � O� 1� and are responsible for the experimentally observed reduction in the flame front wrinkle size in the flow direction [referred to as “filtering” by Bourehla and Baillot (Bourehla, A, and Baillot, F, “Appearance and Stability of a Laminar Conical Premixed Flame Subjected to an Acoustic Perturbation,” Combustion and Flame,

Exploring the effects of nanostructured particles on liquid nitromethane combustion

Abstract: Monopropellants consisting of liquid nitromethane and high-surface-area particles of silicon- and aluminum-based oxides were examined to determine the effects of low concentrations of nanostructured materials on deflagration processes. The combustion rates were characterized by measuring linear burning rates in a large pressure vessel filled with argon. Results showed that nitromethane burning rates may be increased at lower pressures with dilute additions of particles. Increases in nitromethane burning rates of greater than 50 % were found with less than 1.0 wt% of particle addition at a nominal pressure of 5.24 MPa. The particle additions were estimated to have only small effects on the equilibrium flame temperatures, density, viscosity, and specific heat. Nitromethane with particle additives displayed a lower burning-rate equation pressure exponent (i.e., reduced pressure sensitivity), which was inversely proportional to particle concentration. Above a pressure of approximately 9 MPa, up to the maximum test pressure (∼14 MPa), the particle additives did not affect the nitromethane burning rate. Condensed-phase temperature profiles of the deflagrating fluid, surface tension, and fluid thermal conductivities were measured in order to elucidate some of the mechanisms causing enhanced burning rates at lower pressures.