Showing papers in "Journal of Propulsion and Power in 2006"

Nickel-Based Superalloys for Advanced Turbine Engines: Chemistry, Microstructure and Properties

Abstract: The chemical, physical, and mechanical characteristics of nickel-based superalloys are reviewed with emphasis on the use of this class of materials within turbine engines. The role of major and minor alloying additions in multicomponent commercial cast and wrought superalloys is discussed. Microstructural stability and phases observed during processing and in subsequent elevated-temperature service are summarized. Processing paths and recent advances in processing are addressed. Mechanical properties and deformation mechanisms are reviewed, including tensile properties, creep, fatigue, and cyclic crack growth. I. Introduction N ICKEL-BASED superalloys are an unusual class of metallic materials with an exceptional combination of hightemperature strength, toughness, and resistance to degradation in corrosive or oxidizing environments. These materials are widely used in aircraft and power-generation turbines, rocket engines, and other challenging environments, including nuclear power and chemical processing plants. Intensive alloy and process development activities during the past few decades have resulted in alloys that can tolerate average temperatures of 1050 ◦ C with occasional excursions (or local hot spots near airfoil tips) to temperatures as high as 1200 ◦ C, 1 which is approximately 90% of the melting point of the material. The underlying aspects of microstructure and composition that result in these exceptional properties are briefly reviewed here. Major classes of superalloys that are utilized in gas-turbine engines and the corresponding processes for their production are outlined along with characteristic mechanical and physical properties.

Gas Turbine Film Cooling

Abstract: The durability of gas turbine engines is strongly dependent on the component temperatures. For the combustor and turbine airfoils and endwalls, film cooling is used extensively to reduce component temperatures. Film cooling is a cooling method used in virtually all of today's aircraft turbine engines and in many power-generation turbine engines and yet has very difficult phenomena to predict. The interaction of jets-in-crossflow, which is representative of film cooling, results in a shear layer that leads to mixing and a decay in the cooling performance along a surface. This interaction is highly dependent on the jet-to-crossflow mass and momentum flux ratios. Film-cooling performance is difficult to predict because of the inherent complex flowfields along the airfoil component surfaces in turbine engines. Film cooling is applied to nearly all of the external surfaces associated with the airfoils that are exposed to the hot combustion gasses such as the leading edges, main bodies, blade tips, and endwalls. In a review of the literature, it was found that there are strong effects of freestream turbulence, surface curvature, and hole shape on the performance of film cooling. Film cooling is reviewed through a discussion of the analyses methodologies, a physical description, and the various influences on film-cooling performance.

Continuous Spin Detonations

Abstract: Results on controlled continuous spin detonation of various fuels in liquid-propellant rocket motors and ramjet combustors are reported. Schemes of chambers, combustion in transverse detonation waves, and typical photographic records of transverse detonation waves are given. The flow structure, existence conditions, and basic properties of continuous detonation are considered. An analysis of physical, chemical, and geometric parameters determining spin detonation is presented. Results of studying continuous spin detonation of C 2 H 2 + air and H 2 + air mixtures in an annular ducted chamber 30.6 cm in diameter are reported. The range of existence of continuous spin detonation in fuel-air mixtures is determined as a function of the governing parameters. In the case of high-quality mixing, the transverse detonation wave velocity and structure are extremely stable in a wide range of the ratios of propellant components and in the examined range of pressures in the chamber.

Modeling and Analysis of Mistuned Bladed Disk Vibration: Status and Emerging Directions

Abstract: The literature on reduced-order modeling, simulation, and analysis of the vibration of bladed disks found in gas-turbine engines is reviewed. Applications to system identification and design are also considered. In selectively surveying the literature, an emphasis is placed on key developments in the last decade that have enabled better prediction and understanding of the forced response of mistuned bladed disks, especially with respect to assessing and mitigating the harmful impact of mistuning on blade vibration, stress increases, and attendant high cycle fatigue. Important developments and emerging directions in this research area are highlighted.

Erosion and Deposition in Turbomachinery

Abstract: This paper presents a review of erosion and deposition research in turbomachines and the associated degradation in engine performance caused by particulate matter ingestion Parameters affecting surface material losses as a result of erosion and development of experimental and analytical approaches to predict flowpath erosion and deposition are discussed Tests results that quantify the effects of temperature, impact particle composition, impact velocity and angle, and surface material composition are reviewed along with particle restitution data (ratios of rebound to impact velocities and angles) Development and application of models using these data to calculate surface erosion in turbomachinery are described These models predict particle trajectories in turbomachinery passages to determine impact rates, impact velocities, impact angles and uses the experimentally-obtained erosion data to calculate material losses Literature on the effects of erosion on turbomachine performance and life is surveyed Mechanisms of particle delivery and attachment upon arrival at turbomachine flowpath surfaces are also discussed along with experiential models that have been developed to predict surface deposit buildup Delivery to turbine surfaces can occur as a result of inertial flight, as for erosion, but also through transport mechanisms involving turbulence, Brownian diffusion, and thermophoresis The particle size range, where each of these mechanisms is dominant for delivery to surfaces, is described The history and experience of developing models that use these mechanisms to quantify particle delivery rates to turbine flow path surfaces is discussed, along with the use of sticking fraction data to determine the amount of material retained on the surfaces after delivery and the resulting deposit buildup rates Finally, factors that control whether extreme rates of deposition can occur in turbomachinery are described

Thermal Barrier Coatings for Aeroengine Applications

Abstract: Thermal barrier coatings (TBCs) have been in use in aeroturbine engine hot sections for over 20 years. The initial applications were driven by the need to suppress component degradation caused by excessive thermal gradients in vane airfoils. A TBC is essentially a layered, multimaterial structure that must withstand harsh temperature, environmental, thermal cycling, and stress conditions for an extended number of aircraft takeoffs and landings. A description of the TBC materials systems presently in use is presented, together with a summary of current understanding of materials and failure issues in TBCs.

Sealing in Turbomachinery

Abstract: Clearance control is of paramount importance to turbomachinery designers and is required to meet today's aggressive power output, efficiency, and operational life goals. Excessive clearances lead to losses in cycle efficiency, flow instabilities, and hot gas ingestion into disk cavities. Insufficient clearances limit coolant flows and cause interface rubbing, overheating downstream components and damaging interfaces, thus limiting component life. Designers have put renewed attention on clearance control, as it is often the most cost effective method to enhance system performance. Advanced concepts and proper material selection continue to play important roles in maintaining interface clearances to enable the system to meet design goals. This work presents an overview of turbomachinery sealing to control clearances. Areas covered include: characteristics of gas and steam turbine sealing applications and environments, benefits of sealing, types of standard static and dynamics seals, advanced seal designs, as well as life and limitations issues.

Axial Turbine Blade Tips: Function, Design, and Durability

Abstract: An overview of the science and technology involved in today's turbine engines is presented with specific focus on the critical rotational-to-stationary interfaces comprising axial turbine blade tips. The purpose is to provide a concise informative review of turbine blade tip functional, design, and durability issues. Neither a historical account nor a bibliography is presented. Attention is paid primarily to the most challenging blade tips in high-pressure, high-temperature gas turbine systems, although most of the science discussed applies equally well to blade tips in low-pressure turbines, as well as steam turbines. As such, a wide range of both aircraft engine and power generating turbine systems are considered. Basic functional requirements, turbine systems design aspects, and transient operational considerations affecting blade tips and affected by blade tips are discussed in light of the multidisciplinary tradeoffs involved in a successful design. The three dominant design philosophies for blade tips in practice today are presented with detailed examination of the aerodynamics, heat transfer, and cooling benefits and detractors. Finally, the in-service durability aspects of turbine blade tips are noted.

Characteristics of Dual Transverse Injection in Scramjet Combustor, Part 1: Mixing

Abstract: The mixing characteristics of a dual transverse injection system in a scramjet combustor were studied with numerical methods. The effects of the jet-to-cross-flow momentum flux ratio and the distance between injectors on mixing characteristics were investigated. Three-dimensional Navier–Stokes equations, including the k–! SST turbulencemodel,weresolvedwiththe finitevolumemethodadoptingtheupwindmethodofEdwards’lowdiffusion flux splitting scheme. It is shown that the mixing characteristics of a dual transverse injection system are very differentfromthoseofasingleinjectionsystem.Therearinjection flowisstronglyinfluencedbyblockageeffectsdue to the momentum flux of the front injection flow and thus has higher expansion and penetration than the front injection flow. The dual injection system has a higher mixing rate and a higher penetration but have more losses of stagnation pressure than the single injection system. It is also shown that there is an optimal distance between injectors for mixing characteristics and that the optimal distance increased as the jet-to-cross-flow momentum flux ratio increased.

Propulsive performance of airbreathing pulse detonation engines

Abstract: The propulsive performance of airbreathing pulse detonation engines at selected flight conditions is evaluated by means of a combined analytical/numerical analysis. The work treats the conservation equations in axisymmetric coordinates and takes into account finite-rate chemistry and variable thermophysical properties for a stoichiometric hydrogen/air mixture. In addition, an analytical model accounting for the state changes of the working fluid in pulse detonation engine operation is established to predict the engine performance in an idealized situation. The system under consideration includes a supersonic inlet, an air manifold, a valve, a detonation tube, and a convergent-divergent nozzle. Both internal and external modes of valve operation are implemented. Detailed flow evolution is explored, and various performance loss mechanisms are identified and quantified. The influences of all known effects (such as valve operation timing, filling fraction of reactants, nozzle configuration, and flight condition) on the engine propulsive performance are investigated systematically. A performance map is established over the flight Mach number of 1.2-3.5. Results indicate that the pulse detonation engine outperforms ramjet engines for all the flight conditions considered herein. The benefits of pulse detonation engines are significant at low-supersonic conditions, but gradually decrease with increasing flight Mach number.

High-specific impulse Hall thrusters, part 1: Influence of current density and magnetic field

Abstract: A laboratory-model Hall thruster with a magnetic circuit designed for high-specific impulse (2000‐3000 s) was evaluated to determine how current density and magnetic field affect thruster operation. Results have shown for the first time that a minimum current density and optimum magnetic field shape exist at which efficiency will monotonically increase with specific impulse. At the nominal mass flow rate of 10 mg/s and between discharge voltages of 300 and 1000 V, total specific impulse and total efficiency ranged from 1600 to 3400 s and 51 to 61%, respectively. Comparison with a similar thruster showed how efficiency can be optimized for specific impulse by varying the shape of the magnetic field. Plume divergence decreased from a maximum of 48 deg at 400 V to a minimum of 35 deg at 1000 V, but increased between 300 and 400 V as the likely result of a large increase in discharge current oscillations. The breathing-mode frequency continuously increased with voltage, from 14.5 kHz at 300 V to 22 kHz at 1000 V, in contrast to other Hall thrusters where a sharp decrease of the breathing-mode frequency was found to coincide with increasing electron current and decreasing efficiency. These findings suggest that efficient, high-specific impulse operation was enabled through the regulation of the electron current with the applied magnetic field.

Investigation of Vaporized Kerosene Injection and Combustion in a Supersonic Model Combustor

Abstract: Injection and combustion of vaporized kerosene was experimentally investigated in a Mach 2.5 model combustor at various fuel temperatures and injection pressures. A unique kerosene heating and delivery system, which can prepare heated kerosene up to 820 K at a pressure of 5.5 MPa with negligible fuel coking, was developed. A three-species surrogate was employed to simulate the thermophysical properties of kerosene. The calculated thermophysical properties of surrogate provided insight into the fuel flow control in experiments. Kerosene jet structures at various preheat temperatures injecting into both quiescent environment and a Mach 2.5 crossflow were characterized. It was shown that the use ofvaporized kerosene injection holds the potential of enhancing fuel-air mixing and promoting overall burning. Supersonic combustion tests further confirmed the preceding conjecture by comparing the combustor performances of supercritical kerosene with those of liquid kerosene and effervescent atomization with hydrogen barbotage. Under the similar flow conditions and overall kerosene equivalence ratios, experimental results illustrated that the combustion efficiency of supercritical kerosene increased approximately 10-15% over that of liquid kerosene, which was comparable to that of effervescent atomization.

Turbine Endwall Aerodynamics and Heat Transfer

Abstract: This review addresses recent literature on turbine passage aerodynamics and endwall heat transfer; articles that describe the endwall flow and cooling problems are summarized, recent activity on improving endwall aerothermal design is discussed, improved cooling schemes are proposed, and methods for managing secondary flows to allow more effective cooling are suggested. Much attention is given to aerodynamic losses associated with secondary flows developed near the endwalls. The endwall region flowfield is influenced by the stagnation zones established as the endwall approach flow boundary layer meets the airfoil leading edges, by the curvature of the passages, by the steps and gaps on the endwall surface ahead of and within the passage, by the leakage and coolant flows introduced through the endwall surface ahead of and within the passage, by the tip leakage flows between the blades and shroud in the rotor endwall region, and by many more effects. Recent combustor redesigns have flattened the turbine inlet temperature profile and have raised the turbine inlet temperatures. This, coupled with a continued need to improve engine durability and availability, has spurred strong interest in thermal control of the turbine endwall regions. Thus, much of the literature presented herein is focused on endwall cooling and, in particular, the effects of near-endwall secondary flows on endwall cooling.

To the question of energy use of detonation combustion

Abstract: D URING detonation combustion1 of explosive gas mixtures, immediately after the passing of detonation wavefront (completion of chemical reaction), the products of combustion are in a state (call it state D) that is quite rich in both heat and kinetic energy (i.e., the energy of translational motion). Assuming no losses, the state D of combustion products can be calculated using the classical thermodynamic theory of Jouguet. The propagation speed of detonation obtained in this calculation agrees well with experiment, which confirms the correctness of the thermodynamic theory as the limiting case with no losses. For the combustion products in the detonation wave, such a calculation gives: density is 2–1.7 times greater than that of the initial mixture (approximately (k + 1)/k times greater, where k is the adiabatic exponent, pvk = const, for the combustion products); pressure is approximately 2 times greater than the pressure achieved in a closedvolume explosion; temperature is 10-20% higher than the temperature of a closed-volume explosion (approximately 2k/(k + 1) times higher); speed of translational motion is about 0.4–0.5 times that of detonation propagation, which means that kinetic energy of translational motion reaches 17% of the total energy of the mixture. It is interesting to consider to what extent detonation combustion of fuel allows more efficient energy use. One often comes across proposals for detonation regime of combustion in machines such as a gas turbine. Below we present a thermodynamic analysis of the efficiency of cycles with detonation. We will consider here the essential aspects and only briefly comment about technical viability of cycles, losses that reduce the efficiency as compared to the calculated values, etc.

High-Specific Impulse Hall Thrusters, Part 2: Efficiency Analysis

Abstract: Performance and plasma measurements of a high-specific impulse (2000‐3000 s) Hall thruster were analyzed using a phenomenological performance model that accounted for a partially ionized plasma containing multiply charged ions. Anode efficiency over discharge voltages of 300‐900 V ranged from 57 to 69%, which corresponded to 89‐97% voltage utilization, 86‐90% mass utilization, 77‐81% current utilization, and 97‐99% charge utilization. Although the net decrease of efficiency due to multiply charged ions was at most 3%, the effects of multiply charged ions on the discharge current could not be neglected because the increase of the discharge current with voltage was primarily due to the increasing fraction of multiply charged ions. This and the fact that the maximum deviation of the electron current from its average value was only +5/−14% illustrated how efficient operation at high-specific impulse was enabled through the regulation of the electron current with the applied magnetic field. The electron Hall parameter, defined by acceleration zone plasma properties, was nearly constant with voltage, decreasing from an average of 210 at 300 V to an average of 160 between 400 to 900 V.

Influence of a Conical Axial Injector on Hybrid Rocket Performance

Abstract: This paper was aimed at analyzing the static engine firings results obtained by means of a hybrid rocket where gaseous oxygen was supplied into axial-symmetric polyethylene cylindrical grains through two different injector configurations: an axial conical subsonic nozzle and a radial injector. The axial injector is considered interesting because of its easy design and the remarkable feature that it produces no significant pressure oscillations for the stabilizing effect due to the hot gas recirculation zone established within the combustion port. To take advantage of its qualities, the assessment of the regression rate variations under the flow field generated by this configuration is required. For the investigated set of operating conditions, the instantaneous regression rates exhibit a time-and-space dependence caused by the impinging jet zone dynamics, while the average regression rates are higher and less mass flux dependent than those achieved with the radial injection motor and expected from the turbulent flow through pipes. A comparison to the data from the radial injector was further drawn in terms of combustion efficiency, fuel consumption profiles, and combustion stability. The radial injector, at the same mass flux and pressure, produces lower regression rates, high pressure oscillations and worse combustion efficiency at the same L*, but more uniform fuel consumption.

Efficiency Analysis of a Hall Thruster Operating with Krypton and Xenon

Abstract: Krypton has recently become the focus of attention in the Hall thruster community because of its relatively large specific impulse compared with xenon and its potential to operate with comparable efficiencies. However, before krypton can be considered a viable propellant choice for missions, the performance gap between xenon and krypton must be reduced. A series of diagnostic measurements are taken for xenon and krypton propellant using the NASA-173Mv1 Hall thruster and the results are analyzed using a phenomenological performance model. The combined use of experiments and modeling enables a direct comparison of several efficiency components for each propellant to be made. With this method, it is possible to pinpoint the exact causes for the efficiency gap between xenon and krypton. It is also possible to see the effect of the magnetic field topology on Hall thruster performance and where gains are being made due to the magnetic field. Although there is a large series of competing components that differentiate krypton and xenon performance, the largest factors that dictate the efficiency difference between krypton and xenon are krypton's inferior propellant utilization and beam divergence. For xenon, the propellant utilization is 5-10% higher and the beam divergence efficiency is approximately 8% higher.

Thermodynamic Cycle Analysis for Propagating Detonations

Abstract: Propagating detonations have recently been the focus of extensive work based on their use in pulse detonation engines [1]. The entropy minimum associated with Chapman–Jouguet (CJ) detonations [2] and its potential implications on the thermal efficiency of these systems [3] has been one of the main motivations for these efforts. The notion of applying thermodynamic cycles to detonation was considered first by Zel’dovich [4], who concluded that the efficiency of the detonation cycle is slightly larger than that of a cycle using constant-volume combustion. More recently, Heiser and Pratt [3] conducted a thermodynamic analysis of the detonation cycle for a perfect gas using a one-γ model of detonations. Other studies have used constant-volume combustion as a surrogate for the detonation process [5]. This work presents two main contributions. First, we present an alternative physical model for the detonation cycle handling propagating detonations in a purely thermodynamic fashion. The Fickett–Jacobs (FJ) cycle is a conceptual thermodynamic cycle that can be used to compute an upper bound to the amount of mechanical work that can be obtained from detonating a given mass of explosive. Second, we present computations of the cycle thermal efficiency for a number of fuel-oxygen and fuel-air mixtures using equilibrium chemistry, and we discuss the strong influence of dissociation reactions on the results.

Transpiration Cooling Experiment for Scramjet Engine Combustion Chamber by High Heat Fluxes

Abstract: A transpiration cooling experiment using an optical heating method that provided a heat flux as high as 234 W/cm 2 on the surface of specimen for a scramjet engine was performed. In this experiment, 19-mm-diam sintered, powdered, stainless-steel tubes and a 25-mm square cross-sectional tube of perforated Inconel tube were used to investigate transpiration cooling effectiveness. The cooling effectiveness by transpiration for each specimen was measured and analyzed. As a result, the transpiration cooling mechanism appeared to be a promising approach to remove a large amount of heat from the engine wall. A preliminary analysis of the transpiration cooling mechanism and a scaling conversion study that translates the results from helium tests into the case when a hydrogen medium is used are included.

Forced Response of a Swirling, Premixed Flame to Flow Disturbances

Abstract: This paper presents measurements of the nonlinear response of a premixed flame to harmonic oscillations. These measurements were obtained to improve understanding of nonlinear flame dynamics in unstable combustors. Simultaneous measurements of CH ∗ and OH ∗ chemiluminescence, pressure, and velocity were obtained over a range of forcing amplitudes and frequencies. These data show that the flame chemiluminescence response to imposed oscillations saturates at pressure and velocity amplitudes on the order of p � /p0 ∼ 0.02 and u � /u0 ∼ 0.3. The value of the fluctuating to the mean chemiluminescence signal at saturation varied with equivalence ratio. The phase between the chemiluminescence and acoustic signal also exhibits substantial amplitude dependence, even at driving levels where their amplitude ratios are nearly linear. In contrast, the pressure-velocity relationship remains linear with constant phase over the entire amplitude range. Thus, these results suggest that the nonlinear dynamics of premixed combustors are controlled by the acoustics-heat-release relationship, as opposed to gas dynamical processes. The mechanism for saturation of the flame response appears to be caused by nonlinear interactions between the flow forcing and the parametric flame instability, possibly through their impact on the fluctuating flame position. This parametric instability occurs at half the forcing frequency and is caused by the oscillating flow acceleration in the presence of the density jump at the flame. To our knowledge, this observation of the parametric instability is the first in a turbulent, swirl-stabilized flame. Interactions between the nonlinear heat release and linear combustor acoustics were characterized by sweeping the driving frequency through a resonant combustor frequency. With increasing driving amplitude, the resulting frequency response curves began to saturate in amplitude and bend over toward lower frequencies. Similar to the classical nonlinear Duffing oscillator, this bending of the frequency response curves manifests itself as a subcritical bifurcation in combustor response, whose gain and phase exhibited jumps and hysteresis. The bifurcation structure was measured in detail over a range of conditions and shown to have a two-dimensional dependence upon both amplitude and frequency.

High-Frequency Transverse Acoustic Coupling in a Multiple-Injector Cryogenic Combustor

Abstract: High-frequency combustion oscillations are investigated experimentally. The combustor fed by cryogenic propellants operates under elevated pressure conditions (p c = 0.9 MPa) and is equipped with three coaxial injectors fed by liquid oxygen and gaseous methane. Injection parameters are in the typical range used in rocket engines. This experiment simulates on a model scale conditions prevailing in such systems, but full similarity is not achieved. The chamber exhibits a set of resonant modes with eigenfrequencies above 1 kHz. The study focuses on high-frequency dynamics resulting from a strong coupling between one of the transverse modes and combustion. The combustor is forced with an external actuator. The eigenmodes are identified with a linear frequency sweep, and then the system is modulated at the first transverse resonant frequency. The flame motion and response are observed with a high speed and two intensified charge-coupled-device cameras recording phase-conditioned images. In a set of experiments carried out on the multiple-injector combustor, operating conditions were changed systematically to determine parameter ranges leading to combustion sensitivity to transverse excitation. Strong coupling is observed in this way with a spectacular modification of the flame spread. Emission from the three flames is notably intensified when this coupling occurs, whereas thermocouples placed on the lateral walls detect a rapid increase in temperature. The OH* emission intensity that can be linked to the heat-release rate is increased. A phase analysis indicates that the pressure and OH* emission oscillate transversally and in phase at the modulation frequency. This behavior is also observed with the high-speed camera, which also features enhanced reactive vortices convected in the downstream direction at a lower frequency.

Liquid-Fueled Detonations in Tubes

Abstract: Motivated by the current interest in pulse detonation engines, a brief review of previous work on liquid-fueled detonations in tubes is presented. Although the characteristics of detonations in high-vapor-pressure liquid fuels are analogous to those of gaseous detonations, the behavior of detonations in low-vapor-pressure liquid fuels is quite different. The multidimensional structure and the cell sizes of such mixtures have not yet been established. Theoretical, computational, and experimental studies have shown that if the droplets are fine enough (less than about 10 μm), detonations propagating at about the equivalent gas-phase Chapman‐Jouguet detonation velocity can be attained in a variety of fuel‐oxygen systems. With larger droplet sizes, a deficit in the propagation velocity compared to the gas-phase velocity is observed. Various explanations for this deficit are presented. The droplet size required for detonations in fuel‐air mixtures is smaller than that in fuel‐oxygen mixtures. Heating the initial mixture helps both by easing the detonation initiation process and by aiding in the detonation propagation by introducing some prevaporization and reducing the droplet size. The need for additional research on multiphase detonations to aid in the further development of liquid-fueled detonation engines is also emphasized.

Experimental and numerical study of a dual-mode scramjet combustor

Abstract: A Mach 2, hydrogen-air combustor with an unswept 10-deg ramp fuel injector was experimentally and numerically studied for a simulated flight Mach number near 5. Numerical modeling was performed using the Viscous Upwind Algorithm for Complex Flow Analysis code, and results were compared against experimental wall-pressure distributions, fuel plume images, and fuel plume velocity measurements. The model matched wall-pressure distributions well for the case of fuel-off and fuel-air mixing. For a fuel-air reacting case, pressure was matched well in the upstream third of the duct. Downstream, however, the pressure rise as a result of combustion was underpredicted. Based on the fuel plume imaging and velocity measurements, fuel plume shape was matched well for both the mixing and reacting cases. However, plume size, penetration, and centerplane axial growth were generally underpredicted by the model. The full extent of the velocity reduction caused by thermal choking was also not predicted. Despite these findings, the numerical model performed better than a previous model developed by the investigators. It was proposed that differences between the present numerical model and experiment stemmed from numerical underprediction of fuel-air turbulent mixing, and this resulted in underprediction of heat release.

Numerical studies on specific impulse of partially filled pulse detonation rocket engines

Abstract: Performance analyses of pulse detonation rocket engines (PDREs) were numerically studied, focusing on partialfill effects at ground tests. The initial detonable mixture, inert gas, fuel-fill fraction, equivalence ratio, and initial temperature of inert gas were changed as governing parameters. The simulation results were compared against those of previous studies and agreed well with them. The simulation results indicated that the initial mass fraction of the detonable mixture to total mass of the gas was the predominant factor for the specific impulse of partially filled PDREs. Based on the numerical results, a new, simple empirical formula is proposed to predict the specific impulse of partially filled PDREs.

Experimental and computational investigation of light-gas injectors in mach 4.0 crossflow

Abstract: An experimental and computational study of an aerodynamic ramp (aeroramp) injector was conducted at Virginia Polytechnic Institute and State University. The aeroramp consisted of an array of two rows with two columns of flush-wall holes that induce vorticity and enhance mixing. The holes were spaced four diameters apart in the streamwise direction with two diameters transverse spacing between them. For comparison, a single-hole circular injector with the same area angled downstream at 30 deg was also examined. Test conditions involved sonic injection of helium heated to 313 K to safely simulate hydrogen into a Mach 4 air cross stream with average Reynolds number 5.77 e+7 per meter at a jet to freestream momentum flux ratio of 2.1. Sampling probe measurements were utilized to determine the local helium concentration. Pitot and cone-static-pressure probes and a diffuser thermocouple probe were employed to document the flow. This allowed total pressure losses to be determined. The numerical flow solver used was GASP v. 4.2. The inviscid fluxes were computed in three dimensions using third-order AUSM+ with modified essentially nonoscillatory limiting. The AUSM+ algorithm was chosen because of its good resolution of shock discontinuities and its efficiency. The Wilcox k-ω (1998) turbulence model was used. The main results of this work can be summarized as follows: 1) the mixing efficiency value of this aeroramp design, which was optimized at Mach 2.4 for hydrocarbon fuel, was only slightly higher than that of the single-hole injector at these flow conditions; 2) the mass-averaged total pressure loss parameter showed that the aeroramp and single-hole injectors had the same overall losses; 3) the computational fluid dynamics (CFD) was unable to accurately predict the quantitative mixing data produced by the experiment, however, the qualitative comparisons of the injectors using the CFD predictions agreed with the experiment.

Faraday acceleration with radio-frequency assisted discharge

Abstract: A new electrodeless accelerator concept that relies on an rf-assisted discharge, an applied magnetic field, and electromagnetic acceleration using an inductive coil is presented. The presence of a preionized plasma allows for current sheet formation at lower discharge voltages and energies than those found in other pulsed inductive accelerator concepts. A proof-of-concept experiment, supported by optical and probe diagnostics, has been constructed and used to demonstrate low-voltage, low-energy current sheet formation and acceleration. Magnetic field data indicate that the peak sheet velocity in this unoptimized configuration operating at a pulse energy of 78.5 J is 12 km/s. Visual observations indicate that plasma follows the applied magnetic field from the rf discharge to the face of the planar acceleration coil, while magnetic field probing and visualization using a fast-framing camera show the formation and acceleration of the current sheet.

Turbine Blade Internal Cooling Passages with Rib Turbulators

Abstract: Gas turbines are extensively used for aircraft propulsion, land-based power generation, and industrial applications. Developments in turbine cooling technology play a critical role in increasing the thermal efficiency and power output of advanced gas turbines. Gas turbine blades are cooled internally by passing the coolant through several rib-enhanced serpentine passages to remove heat conducted from the outside surface. For internal cooling, focus is placed on the effect of rotation on rotor blade coolant passage heat transfer with rib turbulators. In particular, the most recent publications are covered that deal with the rotational effects on internal cooling passage heat transfer with low and high aspect ratio channels with various high-performance rib geometries. To better understand the complex three-dimensional flow physics in the complicated blade internal coolant passage geometry, the computational flow and heat transfer results are presented and compared using the Reynolds averaged Navier-Stokes method with various turbulence models such as k-e, and second-moment closure models.

Simulation of Void Collapse in an Energetic Material, Part I: Inert Case

Abstract: An Eulerian, sharp-interface, fixed Cartesian grid method is applied to study hot-spot formation in an energetic material (HMX) subject to shock loading. The mass, momentum, and energy equations are solved along with evolution equations for deviatoric stresses and equivalent plastic strain. Pressure is obtained from the Mie-Griineisen equation of state. The material is modeled as a viscoplastic solid. High-order accurate essentially-nonoscillatory (ENO) shock-capturing schemes along with a particle-level set technique are used to evolve sharp immersed boundaries. The details of void collapse under shock loading and the resulting conversion of mechanical energy into localized regions of high thermal energy (hot spots) in the solid material are analyzed. Insights into the precise mechanisms of initiation sensitivity as a result of hot-spot formation in porous energetic materials are obtained.

Characteristics of Dual Transverse Injection in Scramjet Combustor, Part 2: Combustion

Abstract: The combustion characteristics of a dual transverse injection system in a scramjet combustor were studied with numerical methods. The effects of the jet-to-crossflow momentum flux ratio and the distance between injectors on combustion characteristics were investigated. It is shown that the dual injection system has very different combustion characteristics with respect to the single injection system; the burning process of the rear injection flow is strongly influenced not only by the blockage effects but also by the preheating effects due to the chemical reactions of the front injection flow. The dual injection system has a higher burning rate and a higher flame height but more loss of stagnation pressure than the single injection system. It is also shown that there is an optimal distance between injectors for combustion characteristics and that the optimal distance increases as the jet-to-crossflow momentum flux ratio increases.

Model for the Performance of Airbreathing Pulse-Detonation Engines

Abstract: A simplified flowpath analysis of a single-tube airbreathing pulse detonation engine is described. The configuration consists of a steady supersonic inlet, a large plenum, a valve, and a straight detonation tube (no exit nozzle). The interaction of the filling process with the detonation is studied, and it is shown how the flow in the plenum is coupled with the flow in the detonation tube. This coupling results in total pressure losses and pressure oscillations in the plenum caused by the unsteadiness of the flow. Moreover, the filling process generates a moving flow into which the detonation has to initiate and propagate. An analytical model is developed for predicting the flow and estimating performance based on an open-system control volume analysis and gasdynamics. The existing single-cycle impulse model is extended to include the effect of filling velocity on detonation tube impulse. Based on this, the engine thrust is found to be the sum of the contributions of detonation tube impulse, momentum, and pressure terms. Performance calculations for pulse detonation engines operating with stoichiometric hydrogen–air and JP10–air are presented and compared to the performance of the ideal ramjet over a range of Mach numbers.