Showing papers in "Journal of Propulsion and Power in 1991"
TL;DR: In this article, a theoretical analysis is made of supersonic flow of a combustible gas mixture past a wedge or an inclined wall, which shows that, for approach velocities roughly 25% or more greater than the Chapman-Jouguet velocity of the reactant mixture, there exists a usefully wide range of turning angles within which ODWs may be attached or stabilized.
Abstract: Proposals have been made to utilize stabilized oblique detonation waves (ODWs) for the propulsion of hypersonic air-breathing vehicles and hypervelocity mass launchers. There exists hypersonic flight regimes where premixing of fuel and air may be desirable or unavoidable due to finite chemical induction times. Consequently, it is essential to understand under what conditions detonations may occur in order to design supersonic combustors to either avoid or utilize them efficiently. A theoretical analysis is made of supersonic flow of a combustible gas mixture past a wedge or an inclined wall, which shows that, for approach velocities roughly 25% or more greater than the Chapman-Jouguet velocity of the reactant mixture, there exists a usefully wide range of turning angles within which ODWs may be attached or stabilized. For smaller wedge angles, either an incomplete ODW, shock-induced combustion, or no combustion at all may ensue. For larger wedge angles, the wave will detach and form an overdriven normal detonation or normal shock-induced combustion wave immediately upstream of the stagnation point, decaying off axis to either a single oblique Chapman-Jouguet wave, or bifurcating to form an oblique shock followed by a shock-induced deflagration.
185 citations
TL;DR: In this article, a thermally choked ram accelerator, a ramjet-in-tube device for accelerating projectiles to ultrahigh velocities, is investigated theoretically and experimentally.
Abstract: Operational characteristics of the thermally choked ram accelerator, a ramjet-in-tube device for accelerating projectiles to ultrahigh velocities, are investigated theoretically and experimentally. The projectile resembles the centerbody of a conventional ramjet and travels through a stationary tube filled with premixed gaseous fuel and oxidizer at high pressure. The combustion process travels with the projectile, its thermal choking producing a pressure field which results in thrust on the projectile. The results of experiments with 45-75 gm projectiles in a 12.2 m long, 38 mm bore accelerator, using methane-based propellant mixtures, are presented in the velocity range of 1150-2350 m/s. Acceleration of projectiles with staged propellants and transitions between different mixtures are investigated and the velocity limits in several propellant mixtures are explored. Agreement between theory and experiment is found to be very good.
144 citations
TL;DR: In this article, the authors present experimental and computational studies of the pulsed detonation engine concept (PDEC) and present results of their recent numerical study of this concept, and demonstrate the use of current advances in numerical simulation for the analysis of PDEC.
Abstract: Here we review experimental and computational studies of the pulsed detonation engine concept (PDEC) and present results of our recent numerical study of this concept. The PDEC was proposed in the early 1940s for small engine applications; however, its potential was never realized due to a complicated, unsteady operation regime. In this study, we demonstrate the use of current advances in numerical simulation for the analysis of the PDEC. The high-thrust/engine volume ratio obtained in our simulations demonstrates promising potential of the pulsed detonation engine concept.
108 citations
TL;DR: A summary of performance and lifetime characteristics of pulsed and steady-state magnetoplasmadynamic (MPD) thrusters is presented in this article, where the technical focus is on cargo vehicle propulsion for exploration-class missions to the moon and Mars.
Abstract: A summary of performance and lifetime characteristics of pulsed and steady-state magnetoplasmadynamic (MPD) thrusters is presented. The technical focus is on cargo vehicle propulsion for exploration-class missions to the moon and Mars. Relatively high MPD thruster efficiencies of 0.43 and 0.69 have been reported at about 5000-s specific impulse using hydrogen and lithium, respectively. Efficiencies of 0.10 to 0.35 in the 1000to 4500-s specific impulse range have been obtained with other propellants (e.g., Ar, NHa, NI). Electrode power losses of less than 20% at megawatt power levels using pulsed thrusters indicate the potential of high MPD thruster performance. Extended tests of pulsed and steady-state MPD thrusters yield total impulses at least two to three orders of magnitude below that necessary for cargo vehicle propulsion. Performance tests and diagnostics for life-limiting mechanisms of megawatt-class thrusters will require high-fidelity test stands, which handle in excess of 10 kA, and a vacuum facility whose operational pressure is less than 4 x 10 ~ Pa.
89 citations
TL;DR: In this paper, the authors review and assess the data base for shock boundary-layer interaction that is pertinent to the flow prediction in supersonic inlets and identify specific areas related to shock wave/boundary layer interaction bleed, for flow separation control.
Abstract: The performance of supersonic inlets is strongly affected by the boundary-layer development over its internal surfaces. Boundary-layer bleed is used to suppress separation and to provide the desired inlet performance. The gain in pressure recovery and stability is accompanied, however, with a loss in mass flow and an increase in drag that must be minimized by optimizing the amount of bleed and bleed configuration. The purpose of this work is to review and assess the data base for shock boundary-layer interaction that is pertinent to the flow prediction in supersonic inlets. The first part of the review concerns mixed compression supersonic inlets and their bleed system performance at design and off-design conditions. Based on the assessment of this data, specific areas related to shock wave/boundary-layer interaction bleed, for flow-separation control, are identified, the last part of the review addresses this phenomena in various twoand three-dimensional flow configurations. The effect of bleed in the interaction zone is especially emphasized.
77 citations
TL;DR: In this paper, the numerical solution of laminar, two-dimensional, compressible, and unsteady Navier-Stokes equations is aimed at a complete description of acoustic boundary layers that develop above a burning pro pell ant.
Abstract: The numerical solution of laminar, two-dimensional, compressible, and unsteady Navier-Stokes equations is aimed at a complete description of acoustic boundary layers that develop above a burning pro pell ant. Such acoustic boundary layers can be responsible for the so-called flow turning losses. They can govern the local unsteady flow conditions that are seen by the burning propellant to which it finally responds. In those respects, a complete understanding of such acoustic boundary layers is essential to improve existing solid rocket stability prediction codes. The full numerical solution of the Navier-Stokes equations incorporates into the analysis all the features of two-dimension al rocket chamber mean flowfield in a natural manner. After a standing wave pattern is established through forcing at a given frequency, a special Fourier treatment is used to transform the numerical results in a form directly comparable to available linear acoustic data. The presented results indicate that the acoustic boundary layer is substantially thinner than predicted by simplified models. Moreover, its acoustic admittance is found to vary significantly along the chamber, a result that is of major importance to stability predictions. Finally, the acoustic field is found to be rotational over a significant volume of the chamber.
73 citations
TL;DR: In this article, non-reacting and combustion tests were performed for subsonic, sonic, and supersonic conditions using noncircular injectors in a gas generator combustor.
Abstract: Nonreacting and combustion tests were performed for subsonic, sonic, and supersonic conditions using noncircular injectors in a gas generator combustor. The noncircular injectors, including square, equilateral-, and isosceles-triangular nozzles, were compared to a circular injector. The flowfields of the jets were mapped with hot-wire anemometry and visualized using spark schlieren photography. The combustion characteristics were visualized by high-speed photography and thermal imaging, and the temperature distribution was measured by a rake of thermocouples . The present tests conducted at high Reynolds and Mach numbers confirmed earlier results obtained for the low range of these numbers, i.e., the combination of large-scale mixing at the flat sides with the fine-scale mixing at the vertices is beneficial for combustion. Large-scale structures provide bulk mixing between the fuel and air, whereas fine-scale mixing contributes to the reaction rate and to better flameholding characteristics.
68 citations
TL;DR: In this article, the axial vorticity mixing mechanism previously shown to be responsible for rapid mixing in low-speed, subsonic flows is also effective in a supersonic flow environment.
Abstract: An experimental study of supersonic mixer nozzles in a coflowing stream has been conducted in the United Technologies Research Center open jet acoustic wind tunnel. Enhanced supersonic jet mixing is important in a number of applications including jet exhaust noise reduction and improved flow distribution within engine combustors. Recently discovered novel concepts promoting enhanced mixing via the introduction of axial vorticity into the exhaust have resulted in studies of the mixing process for nozzles operating at low, subsonic Mach number conditions and low temperatures. The goal of the present experimental study was to evaluate these approaches to jet mixing in the high-temperature, supersonic primary flow regime typical of turbofan/turbojet engine operation. Jet total temperature, total pressure, static pressure, and velocity distributions were measured to characterize the mixing process for baseline slot and circular nozzles, and for several mixer nozzles. The measurements were made at a jet exit Mach number of 1.5, a wind-tunnel forward flight Mach number of 0.5, and a jet total temperature of 1000°F. A principal conclusion of this study is that the axial vorticity mixing mechanism previously shown to be responsible for rapid mixing in low-speed, subsonic flows is also effective in a supersonic flow environment. Reductions in nozzle potential core length of approximately a factor of two relative to the slot nozzle configuration were observed for one of the mixer nozzles studied.
67 citations
TL;DR: In this paper, the behavior of individual boron particles in the flowfield of a solid fuel ramjet (SFRJ) combustor is investigated, and the results demonstrate the limited ranges of particle size and ejection velocity which enable ignition and sustained combustion.
Abstract: Theoretical investigation on the behavior of individual boron particles in the flowfield of a solid fuel ramjet (SFRJ) combustor is presented. The study was motivated by the observed difficulties in achieving good combustion efficiencies of boron required to exploit its remarkable theoretical energetic performance. The equations describing the gas flowfield and the particle behavior are solved numerically. The solution presents the trajectory, temperature, and history of the boron particles due to the interactions with the surrounding gas, as well as the ignition envelope and combustion time. The results demonstrate the limited ranges of particle size and ejection velocity which enable ignition and sustained combustion, reveal why practical systems often exhibit poor combustion efficiencies, and predict the conditions where ignition and efficient combustion of boron are feasible. Nomenclature
64 citations
TL;DR: In this paper, a discrete operating conditions gas path analysis (DOCGPA) algorithm is used to diagnose component directed fault in a commercial turbofan engine using an engine computer model.
Abstract: A common feature of all Differential Gas Path Analysis methods is the necessity of measuring a number of performance variables greater or at least equal to the number of diagnostic parameters which have to be estimated. Discrete Operating Conditions Gas Path Analysis (DOCGPA) is an extended version of the conventional GPA algorithms, providing-among other things-the capability to overcome this problem. In the present paper, we describe how this method can be coupled with an engine computer model, in order to perform component directed fault diagnosis. Application to a commercial turbofan engine demonstrates the effectiveness of the proposed method. Nomenclature [C] = influence coefficient matrix EPR = engine pressure ratio / = scalar defined in Eq. (8) k = number of different operating points [M] = information matrix m = number of unknown component parameters n = number of measured variables [P] = covariance matrix of unknown components parameters m = number of unknown component parameters PEUI = performance estimation uncertainty index (equal to J) [R] = covariance matrix of measures variables tr{ } =' trace of a matrix u = operating condition vector x = component parameter vector y = measurement vector A( ) = percentage change from an initial value 2 = summation Superscripts T = transpose matrix -1 = inverse matrix
60 citations
TL;DR: In this article, an investigation has been conducted of the velocity coupling phenomenon reported in acoustically unstable solid-propellant rocket motors, where dry ice is introduced over the dry-ice surface by means of a mechanically driven piston.
Abstract: An investigation has been conducted of the velocity coupling phenomenon reported in acoustically unstable solid-propellant rocket motors. An innovative simulation facility has been built using solid carbon dioxide as the simulated propellant. Acoustic disturbances are introduced over the dry-ice surface by means of a mechanically driven piston. Experiments have been conducted with dry ice located near both a velocity antinode and a velocity node. Mass flow rate and acoustic pressure measurements indicate the existence of a coupling mechanism, strongly dependent on the acoustic velocity amplitude, between the acoustic disturbance and the dry-ice sublimation process. Flow visualization and hot-film anemometry both show that the flow is turbulent near resonance. Transition to turbulence near a velocity node appears to occur at a smaller critical acoustic velocity amplitude than that near a velocity antinode. Some preturbulent instability has also been observed. Acoustically induced, turbulent forced convection is believed to be responsible for the increase in the sublimation rate of the dry ice (simulated burning of the propellant). Turbulence is believed to be one of the principal mechanisms in the velocity coupling phenomenon. An empirical correlation was developed which appeared to apply to the real propellant cases.
TL;DR: In this paper, a computer algorithm is introduced which can be used to simulate the fluid behavior in that environment, in particular the excitation of sloshing waves due to different gravity environments and rotation speeds.
Abstract: Time-dependent computations have been performed to investigate the dynamical behaviors of fluid under a microgravity environment. A computer algorithm is introduced which can be used to simulate the fluid behavior in that environment, in particular the excitation of sloshing waves due to different gravity environments and rotation speeds. A suggestion on the proper handling and managing of cryogenic fluid propellant to be used in the Gravity Probe-B spacecraft propulsion is made.
TL;DR: In this article, the main combustion characteristics of small-scale solid fuel ramjet combustors were characterized and the specific effects resulting from the special size range of the combustor were pointed out.
Abstract: The objective of this research was to characterize the combustion and flameholding limits in solid fuel ramjet combustors of particularly small size (10mm i.d.). The main combustion characteristics, as well as the specific effects resulting from the special size range, are pointed out. Several hundred static tests were performed with three different fuel types, polymethylmethacrylate (plexiglass), polyethylene, and polybutadiene. Air inlet temperatures of 800 K, 520 K, and room temperature were tested to simulate flight Mach numbers of 3, 2, and low subsonic, respectively. All three fuels exhibited very good flameholding capability at the higher air temperature (800 K). The flammability limits decreased at 520 K and became very narrow at room temperature. The flameholding capability also deteriorated with decreasing combustor size. The flow-reattachment distance, coinciding with the maximum fuel regression rate zone, was shown to be proportional to the inlet step height and was independent of fuel type and port-flow Reynolds number. Nomenclature A — area d = diameter F = thrust G = mass flux H = step height k = thermal conductivity M = Mach number m = exponent m = mass flow rate n = exponent p = pressure Q = heat capacity (sensible enthalpy) Q = heat loss rate q = heat flux Re = Reynolds number r = fuel regression rate T — temperature V = volume xr = reattachment distance Subscripts
TL;DR: In this paper, an experimental method was used to find the rebound characteristics of small solid particles impacting different materials, such data were used for particle trajectories and erosion calculations in turbomachinery.
Abstract: This paper describes an experimental method used to find the rebound characteristics of small solid particles impacting different materials. Such data are used for particle trajectories and erosion calculations in turbomachinery. The materials which are investigated are: 410 Stainless Steel, 2024 Aluminum, 6A1-4V Titanium, INCO 718, RENE 41, AM355, L605 Cobalt, and Alumina (A12O3). Particle materials are fly ash of 15 p. In addition, erosion rate equation was developed for INCO 718 that predicts the erosion rates at high temperatures very well.
TL;DR: In this paper, boron/poly(BA MO/NMMO) fuel-rich solid propellants have been studied due to their potential application to solid-fuel ramjets (SFRJ).
Abstract: Combustion characteristics of boron/poly(BA MO/NMMO) fuel-rich solid propellants have been studied due to their potential application to solid-fuel ramjets (SFRJ). For the same boron content, BAMO/NMMO copolymer-based fuels are superior to conventional hydroxyl terminated polybutadien (HTPB) fuels due to their vigorous pyrolysis characteristic for dispersing boron particles into the main reaction zone. However, their specific impulses are generally lower than that of HTPB, in spite of high positive heats of formation. Formation of hexagonal crystalline boron nitride (BN) has been found in the combustion of this family of propellants studied. Favorable conditions for the formation of BN have been identified. BN also has a significant effect on the theoretical performance for high equivalence ratio conditions. The burning rate was found to depend strongly upon pressure and nonmonotonically on boron content. An "energy sink" hypothesis is proposed to explain this observation. Fine-wire thermocouple measurements support this hypothesis.
TL;DR: In this paper, the properties of the multiphase mixing layer that surrounds the liquid core during atomization breakup were studied, and the results showed that large irregularly shaped liquid elements and drops increased and the proportion of spherical drops decreased with increasing radial distance.
Abstract: The dense-spray region of pressure-atom ized nonevaporating sprays was studied, emphasizing the properties of the multiphase mixing layer that surrounds the liquid core during atomization breakup. The dispersed-phase properties of a large-scale (9.5-mm injector diameter) water jet injected vertically downward in still air were measured using single- and double-pulse holography for both fully developed and slug flow jet exit conditions. The inner portion of the mixing layer contained large irregularly shaped liquid elements and drops, and the proportion of spherical drops increased and drop sizes decreased with increasing radial distance. For present test conditions, the liquid core and the large liquid elements cause mean liquid volume fractions to be high near the axis; however, the gas-containing region was relatively dilute at each instant. Additionally, the velocities of large drops were generally much larger than small drops and predictions based on the locally homogeneous flow approximation, providing direct evidence of significant separated-flow effects in the flow. Finally, the degree of flow development at the jet exit had a substantial effect on the structure of the mixing layer, with increased turbulence levels increasing the number and size of large irregular liquid elements through distortion of the surface of the liquid core—enhancing rates of removal of liquid from the core.
TL;DR: The state of the art in the family of radio-frequency ion thrusters called RIT is discussed in this article, where the R D programs involved in these thrusters are reviewed, and the mode of operation of the RIT thrusters is described.
Abstract: The state of the art in the family of radio-frequency ion thrusters called RIT is discussed. The R D programs involved in these thrusters are reviewed, and the mode of operation of the RIT thrusters is described. The thruster hardware is examined, including the North-South Stationkeeping thrusters RIT 10 and RIT 15 and the primary propulsion thruster RIT 35. 22 refs.
TL;DR: In this paper, the nature and mechanisms of explosions caused by the contact of hypergolic liquid propellants were investigated in detail for several combinations of fuels and oxidizers, and it was shown that the explosion phenomena observed can be classified into three categories.
Abstract: The nature and mechanisms of explosions caused by the contact of hypergolic liquid propellants were investigated in detail for several combinations of fuels and oxidizers. It is shown that the explosion phenomena observed can be classified into three categories. 1) In the case of N2H4/NTO, sudden gasification of a superheated liquid layer formed at the boundary of two liquids occurs spontaneously and a detonation-like reaction proceeds in the reactive mixture produced. 2) In the cases of MMH/NTO and UDMH/NTO, the sudden gasification is caused by the shock of a local ignition, and a turbulent-combustion reaction proceeds in the reactive mixture produced. 3) In the cases of hydrazine type fuels/FNA, the sudden gasification occurs spontaneously as in the case of N2H4/NTO, but it is not augmented by chemical reaction, and in these cases the observed explosion is weak. Information on the vapor layer, which is formed between reactive fuel droplet and pool liquid plays an important role for the occurrence of explosion, is also given, that is based on the high-speed motion picture records.
TL;DR: In this article, an analysis of self-field accelerated quasi-one-dimensional plasma flows with zero axial current is presented, where magnetic convection dominates over magnetic diffusion (large magnetic Reynolds number Rm\ inlet and exit current concentration layers occur, which are analyzed by asymptotic methods to zeroth order in l/Rm.
Abstract: An analysis is presented of self-field accelerated quasi-one-dimensional plasma flows with zero axial current. When magnetic convection dominates over magnetic diffusion (large magnetic Reynolds number Rm\ inlet and exit current concentration layers occur, which are analyzed by asymptotic methods to zeroth order in l/Rm. Sonic passage is found in the inlet layer, which largely determines injector conditions. Near magnetosonic conditions prevail at throats or constant-area sections. Performance factors accounting for inlet convergence and nozzle divergence are derived by exploiting the smallness of pressure forces. The effect of finite magnetic Reynolds numbers is seen to be moderate down to Rm 3 for contoured channels, but to Rm s 10 for constant-area channels. Contouring is shown to be effective for control of current and dissipation concentrations.
TL;DR: In this article, a numerical method for solving the Reynolds-averaged Navier-Stokes equations around axisymmetric afterbody/nozzle configurations, with sharp trailing edges or finite bases, is presented.
Abstract: A numerical method for solving the Reynolds-averaged Navier-Stokes equations around axisymmetric afterbody/nozzle configurations, with sharp trailing edges or finite bases, is presented. Turbulence closure is achieved through either a simple algebraic turbulence model or a low Reynolds number form of the A>e two-equation differential model. The solution procedure uses an explicit time-marching finite-volume method. The performance of each of the turbulence models is assessed through comparisons with experimental data on three series of geometries, including both attached and separated flow cases.
TL;DR: In this paper, a steady and unsteady aerodynamic analysis of ducted fans was developed using a frequency domain panel method based on three-dimensional linear compressible lifting surface theory, where the duct is assumed to be a finite-length right-circular cylinder concentric with the rotor.
Abstract: A steady and unsteady aerodynamic analysis of ducted fans has been developed using a frequency domain panel method based on three-dimensional linear compressible lifting surface theory. The duct is assumed to be a finite-length right-circular cylinder concentric with the rotor. Both the duct and rotor blades are modeled by simple harmonic rotating doublet sheets. The model spans a single reference passage with the influence of the rest of the configuration included by symmetry. Results for the steady state performance characteristics of a ducted rotor are compared with an Euler calculation. The effect of the duct on the unsteady aerodynamic forces induced by blade vibration is examined, and comparisons are made with two-dimensional unsteady cascade theory* Finally, it is shown that the duct has an adverse effect on the aeroelastic stability of the rotor.
TL;DR: In this paper, a detailed study of the two-phase flow produced by a gas-turbine air-blast atomizer is performed with the goal of identifying the interaction between the two phases for both nonreacting and reacting conditions.
Abstract: A detailed study of the two-phase flow produced by a gas-turbine air-blast atomizer is performed with the goal of identifying the interaction between the two phases for both nonreacting and reacting conditions. A two-component phase Doppler interferometry is utilized to characterize three flowfields produced by the atomizer: (1) the single-phase flow, (2) the two-phase nonreacting spray, and (3) the two-phase reacting spray. Measurements of the mean and fluctuating axial and azimuthal velocities for each phase are obtained. In addition, the droplet size distribution, volume flux, and concentration are measured. The results reveal the strong influence of the dispersed phase on the gas, and the influence of reaction on both the gas and the droplet field. The presence of the spray significantly alters the inlet condition of the atomizer. With this alteration quantified, it is possible to deduce that the inertia associated with the dispersed phase damps the fluctuating velocities of the gas. Reaction reduces the volume flux of the droplets, broadens the local volume distribution of the droplets in the region of the reaction zone, increases the axial velocities and radial spread of the gas, and increases the anisotropy in the region of the reaction zone.
TL;DR: In this article, a windowed, two-dimensional solid-fuel-ramjet motor was used with high-speed motion picture cameras to study the effects of fuel composition, pressure, and air mass flux on the surface behavior of metallized fuels within both the recirculation and boundary-layer combustion regions.
Abstract: A windowed, two-dimensional solid-fuel-ramjet motor was used with high-speed motion picture cameras to study the effects of fuel composition, pressure, and air mass flux on the surface behavior of metallized fuels within both the recirculation and boundary-layer combustion regions. Surface behavior and near-surface combustion characteristics were examined to help explain the regression rate/performance characteristics observed in actual motor hardware. Fuels containing no combustion catalyst exhibited the characteristic of extending the combustion envelope to lower pressures with lower values of Shore A hardness. Ignition/flammability limits appeared to be a strong function of surface pyrolysis within the recirculation region. Most metallized fuels exhibited shedding of flakes of (unburned) material from the fuel surface. Flake thickness for boron fuels was typically 200 /*, independent of pressure, and surface areas varied from less than 1 mm2 to approximately 22 mm2. The mass losses attributable to the flaking process play a major role in the overall fuel mass loss mechanism. Bimetallic fuels and fuels containing a combustion catalyst apparently had surface reactions which increase the surface temperature. This should be beneficial for obtaining more complete combustion of the metals within the motor.
TL;DR: In this paper, a combination of tangential and normal air injection into a Mach 3 airflow was experimentally studied, and it was shown that the mixing rate can be significantly increased by the combined tangential-normal injection design over tangential slot injection alone, with up to 92% more entrained mass.
Abstract: A combination of tangential and normal air injection into a Mach 3 airflow was experimentally studied. A rearward facing slot producing tangential injection at a nominal Mach number 1.7 was operated at several different total pressures. An array of transverse tubes of height equal to the slot height and placed just downstream of the slot was operated at two dynamic pressure ratios at both Mach 1 and 2.2. Mean flow measurements of static and total pressures were taken up to 20 slot heights downstream from which Mach number, density and velocity profiles, and entrainment rates were calculated. Various dimensions and spreading angles of the mixing regions were measured directly from nanoshadowgraphs and spark schlieren photographs. Large eddy structures were produced in several cases, leading to increased entrainment of the freestream. For some cases, heated air was injected through the normal tubes, and the jet total temperature decay was measured downstream. It can be seen from the data that the mixing rate can be significantly increased by the combined tangential-normal injection design over tangential slot injection alone, with up to 92% more entrained mass.
TL;DR: In this paper, the two-phase flowfield in the aft-dome region of a solid rocket motor (SRM) with submerged nozzle has been simulated using a combined Eulerian-Lagrangian analysis.
Abstract: The two-phase flowfield in the aft-dome region of a solid rocket motor (SRM) with submerged nozzle has been simulated using a combined Eulerian-Lagrangian analysis. This analysis uses the numerical solution of ensemble averaged Navier-Stokes equations for the continuous (gas) phase coupled with a Lagrangian analysis for the discrete (particulate) phase to simulate the two-phase internal flow. A linearized block-implicit (LBI) scheme is used to solve the governing equations for the continuous phase, which allows the use of a highly stretched grid with sublayer resolution. The motion of the particles is tracked in computational coordinate space resulting in computational efficiency, and the interphase coupling terms for the Eulerian analysis are computed from the instantaneous distribution of the particles. A low Reynolds number form of the k-e turbulence model is used with modifications for injection driven flows. Calculations have been performed for a particular grain configuration of the Space Shuttle SRM. The flowfield in the vicinity of the submerged nozzle, the particle trajectories, and the sensitivity of two-phase effects (such as slag accumulation) to the particle injection parameters are presented in this paper.
TL;DR: In this paper, a two-component laser Doppler velocimeter was used to obtain measurements of the three velocity components and numerous fundamental turbulence quantities in two series of tests with minimum disturbance to the combustor flowfield.
Abstract: Experimental and theoretical studies of nonreacting swirling flow have been performed in a model of an axisymmetric dump combustor. A two-component laser Doppler velocimeter was used to obtain measurements of the three velocity components and numerous fundamental turbulence quantities in two series of tests with minimum disturbance to the combustor flowfield. The results showed the significant effects of swirl, with and without vortex breakdown, on the mean and turbulent flow fields. The experimental results were used to check the performance of a recently developed computer program which used the A>8 closure model. Comparisons of the numerical and experimental results showed the inadequacy of the £-8 turbulence model in representing the complex structure of confined swirling flows.
TL;DR: In this paper, the cooling requirement of a hydrogen-fueled airframe-integrated scramjet engine as well as an airframe was examined, and effects of various parameters including flight Mach number, flight dynamic pressure, engine wall temperature, and engine scale, on the engine characteristics were analyzed.
Abstract: In a previous report, scramjet engine characteristics of different propellant-fed cycles were compared and engine performances were discussed. In this study, the cooling requirement of a hydrogen-fueled airframe-integrated scramjet engine as well as an airframe was examined, and effects of various parameters including flight Mach number, flight dynamic pressure, engine wall temperature, and engine scale, on the engine characteristics were analyzed. The coolant required for the airframe was about 20% of the total coolant. Simple equations that correlate coolant flow rate with those parameters were derived. A B b CD Cp 7sp t M m P Q q
TL;DR: In this article, a design approach was formulated to guide the development effort of the combustor, which combines the capabilities of the analytical tools with well-established empirical correlations to achieve the required performance goals and structural durability of gas turbine combustion systems.
Abstract: To achieve the required performance goals and structural durability of gas turbine combustion systems, a design approach was formulated to guide the development effort of the combustor. The approach combines the capabilities of the analytical tools with well-established empirical correlations. By this means, the impact of systematic modification to the details of the burner is easily determined. The validation effort of the developed model included the utilization of the data obtained for a number of production combustors that significantly varied in design and concept. Fuels used in these combustors included typical aviation fuels such as JP-4 and DF-2 as well as specially prepared high-density fuels. Model validation also involved the application of the model in the development phases of an annular combustor. Several modifications to the dome and primary zone features were proposed in this effort. The predictions of the combustor performance and wall temperatures made using the present approach were found to be in good agreement with the measurements.
TL;DR: In this paper, a one-dimensiona l, stage-by-stag e axial compression system mathematical model has been constructed that can describe system behavior during poststall events such as surge and rotating stall.
Abstract: A one-dimensiona l, stage-by-stag e axial compression system mathematical model has been constructed that can describe system behavior during poststall events such as surge and rotating stall. The model uses a numerical technique to solve the nonlinear, compressible conservation equations of mass, momentum, and energy. Inputs for blade forces and shaft work are provided by a set of quasi-steady stage characteristics modified by a firstorder lagging equation to simulate dynamic stage characteristics. The model was operationally verified using experimental results for a three-stage, low-speed compressor. Using the model, two studies were conducted: one to determine the effect of heat transfer due to rapid system transients on poststall system behavior, and the other to determine the effect of a possible design modification on overall system behavior. Results from these studies demonstrate the use of this modeling technique in studies of compression system poststall behavior.
TL;DR: In this article, the complex three-dimensional flowfield produced by secondary injection of hot gases in a rocket nozzle for thrust vector control is analyzed by solving unsteady three dimensional Euler equations with appropriate boundary conditions, and various system performance parameters like secondary jet amplification factor and axial thrust augmentation are deduced by integrating the nozzle wall pressure distributions obtained as part of the flowfield solution.
Abstract: The complex three-dimensional flowfield produced by secondary injection of hot gases in a rocket nozzle for thrust vector control is analyzed by solving unsteady three-dimensional Euler equations with appropriate boundary conditions. Various system performance parameters like secondary jet amplification factor and axial thrust augmentation are deduced by integrating the nozzle wall pressure distributions obtained as part of the flowfield solution and compared with measurements taken in actual static tests. The agreement is good within the practical range of secondary injectant flow rates for thrust vector control applications.