Showing papers in "Journal of Propulsion and Power in 2004"
TL;DR: A general review of the worldwide evolution of ramjet propulsion since the Wright brothers first turned man's imagination into a practical reality is presented in this article, where the development history and principal contributing development programs are reviewed.
Abstract: A general review is presented of the worldwide evolution of ramjet propulsion since the Wright brothers e rst turned man’ s imagination to e y into a practical reality. A perspective of the technological developments from subsonic to hypersonic e ight speeds is provided to allow an appreciation for the advances made internationally from the early 1900s to current times. Ramjet, scramjet, and mixed-cycle engine types, and their operation and rationale for use are considered. The development history and principal contributing development programs are reviewed. Major airbreathing technologies that had signie cant impact on the maturation of ramjet propulsion and enabled engine designs to mature to their current state are identie ed. The general state of e ight-demonstrated technology is summarized and compared with the technology base of 1980. The current status of ramjet/scramjet technology is identie ed. Ramjet and scramjet propulsion technology has matured dramatically over the years in support of both military and space access applications, yet many opportunities remain to challenge future generations of explorers.
481 citations
TL;DR: In this article, an experimental investigation of the mixing and combustion processes that occur in and around a cavity-based flameholder in a supersonic flow is reported, which is part of an ongoing research program aimed at providing information to help fill these voids and improve the overall understanding of cavities for use as scramjet flameholders.
Abstract: An experimental investigation of the mixing and combustion processes that occur in and around a cavity-based flameholder in a supersonic flow is reported. Cavity-based flameholders are commonly found in hydrocarbon-fueledscramjet combustors; however, detailed information concerning the behavior of these devices, their optimal shape and fueling strategies, combustion stability, and interactions with disturbances in the main airflow (i.e., shock trains or shock-boundary layer interactions) is largely unavailable in the existing literature. This work is part of an ongoing research program aimed at providing information to help fill these voids and improve the overall understanding of cavities for use as scramjet flameholders.
332 citations
TL;DR: In this paper, a class of paraffin-based fuels that burn at surface regression rates that are three to four times that of conventional hybrid fuels is identified, which is a natural attribute of the fuel material, and the use of oxidizing additives or other regression rate enhancement schemes is not required.
Abstract: Recent research at Stanford University has led to the identification of a class of paraffin-based fuels that burn at surface regression rates that are three to four times that of conventional hybrid fuels. The approach involves the use of materials that form a thin, hydrodynamically unstable liquid layer on the melting surface of the fuel. Entrainment of droplets from the liquid-gas interface substantially increases the rate of fuel mass transfer, leading to much higher surface regression rates than can be achieved with conventional polymeric fuels. Thus, high regression rate is a natural attribute of the fuel material, and the use of oxidizing additives or other regression rate enhancement schemes is not required. The high regression rate hybrid removes the need for a complex multiport grain, and most applications up to large boosters can be designed with a single port configuration. The fuel contains no toxic or hazardous components and can be shipped by commercial freight as a nonhazardous commodity. At the present time, grains up to 0.19 m [19.1 cm (7.5 in.)] in diameter and 1.14 m [114.8 cm (45.2 in.)] long are fabricated in a general-purpose laboratory at Stanford University. To further demonstrate the feasibility of this approach, a series of scale-up tests with gaseous oxygen have been carried out using a new Hybrid Combustion Facility (HCF) at NASA Ames Research Center. Data from these tests are in agreement with the small-scale, low-pressure, and low mass flux laboratory tests at Stanford University and confirm the high regression rate behavior of the fuels at chamber pressures and mass fluxes representative of commercial applications.
283 citations
TL;DR: Chinzei et al. as mentioned in this paper investigated the effects of injector geometry on Scramjet Combustor performance and found that injector geometrical geometry has a negative effect on ScRAMJET performance.
Abstract: 3Heiser, W H, Pratt, D T, with Daley, D H, and Mehta, U B, “Hypersonic Airbreathing Propulsion,” AIAA Education Series, AIAA, Washington, DC, 1994, pp 334–370 4Chinzei, N, Komuro, T, Kudo, K, Murakami, A, Tani, K, Masuya, G, and Wakamatsu, Y, “Effects of Injector Geometry on Scramjet Combustor Performance,” Journal of Propulsion and Power, Vol 9, No 1, 1993, pp 146–152 5Chinzei, N, Masuya, G, Kudo, K, Murakami, A, and Komuro, T, “Experiment on Multiple Fuel Supplies to Airbreathing Rocket Combustors,” Journal of Propulsion and Power, Vol 3, No 1, 1987, pp 26–32
235 citations
TL;DR: Choueiri et al. as discussed by the authors idealize the continuous history of the field as a series of five essentially consecutive eras, and define thematic periods in the often continuous stream of events under review and label them as eras.
Abstract: ∗Chair of AIAA’s Electric Propulsion Technical Committee, 2002-2004. Associate Fellow AIAA. Chief Scientist at Princeton University’s Electric Propulsion and Plasma Dynamics Laboratory (EPPDyL). Associate Professor, Applied Physics Group, MAE Department. e-mail: choueiri@princeton.edu. †Presented at the 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Ft. Lauderdale, FL. Copyright c © 2004 by the author. Published by the AIAA with permission. Also published in the Journal of Propulsion and Power, Vol. 20, No. 2, pp. 193–203, MarchApril 2004. When writing history, it is tempting to identify thematic periods in the often continuous stream of events under review and label them as “eras”, or to point to certain achievements and call them “milestones”. Keeping in mind that such demarcations and designations inevitably entail some arbitrariness, we shall not resist this temptation. Indeed, the history of electric propulsion (EP), which now spans almost a full century, particularly lends itself to a subdivision that epitomizes the progress of the field from its start as the dream realm of a few visionaries, to its transformation into the concern of large corporations. We shall therefore idealize the continuous history of the field as a series of five essentially consecutive eras:
190 citations
TL;DR: In this article, a two-dimensional computational model is developed to calculate the heat and mass transport associated with a flowing fuel using a unique global chemical kinetics model, which is validated by measured experimental data obtained from a flow reactor in which ndecane and n-dodecane are mildly cracked.
Abstract: Hydrocarbon fuels have been used as a cooling media in aircraft jet engines for many years. However, advanced aircraft engines are reaching a practical heat transfer limit beyond which the sensible heat transfer provided by fuels is no longer adequate. One solution is to use an endothermic fuel that absorbs heat through chemical reactions. A two-dimensional computational model is developed to calculate the heat and mass transport associated with a flowing fuel using a unique global chemical kinetics model. Most past models do not account for changes in the chemical composition of a flowing fuel and also do not adequately predict flow properties in the supercritical regime. The two-dimensional computational model presented calculates the changing flow properties of a supercritical reacting fuel by use of experimentally derived proportional product distributions. The calculations are validated by measured experimental data obtained from a flow reactor in which n-decane and n-dodecane are mildly cracked. It is believed that these simulations will assist the fundamental understanding of high-temperature fuel flow experiments.
165 citations
TL;DR: In this paper, a multi-objective evolutionary algorithm was used for handling the optimization problem that makes use of Pareto optimality concepts and implements a novel genetic diversity evaluation method to establish a criterion for fitness assignment.
Abstract: A method for transonic compressor multi-objective design optimization was developed and applied to the NASA rotor 37, a test case representative of complex three-dimensional viscous flow structures in transonic bladings. The optimization problem considered was to maximize the isentropic efficiency of the rotor and to maximize its pressure ratio at the design point, using a constraint on the mass flow rate. The three-dimensional Navier‐Stokes code CFXTASCflow ® was used for the aerodynamic analysis of blade designs. The capability of the code was validated by comparing the computed results to experimental data available in the open literature from probe traverses upand downstream of the rotor. A multi-objective evolutionary algorithm was used for handling the optimization problem that makes use of Pareto optimality concepts and implements a novel genetic diversity evaluation method to establish a criterion for fitness assignment. The optimal rotor configurations, which correspond to the maximum pressure ratio and maximum efficiency, were obtained and compared to the original design.
145 citations
TL;DR: In this paper, the combination of aluminum and water was theoretically analyzed to assess its performance potential for space propulsion, in particular for microrocket applications and whenever a compact package is desirable.
Abstract: The combination of aluminum and water was theoretically analyzed to assess its performance potential for space propulsion, in particular for microrocket applications and whenever a compact package is desirable. Heat of reaction, impulse density, and handling safety are features making this combination interesting for chemical thrusters, especially because thrust is higher than typical of satellite electric thrusters. Ideal specific impulse I s p , thrust coefficient, adiabatic flame temperature, and combustion products were calculated for chamber pressures 1-10 atm, nozzle area ratios 25-100, and mixture ratios O/F 0.4-8.0. I s p reaches up to 3500 m/s. Also, the effect of hydrogen peroxide addition to aluminum and water on performance was explored. This combination improves performance slightly at the expense of simplicity, making it less attractive for microrocket engines. Ignition delay times were conservatively estimated assuming that aluminum was coated with its oxide and ignition occurred after the melting of the aluminum oxide. For this purpose heating and kinetics times were evaluated, the first by a one-dimensional physical model, the second by a reduced scheme. Results indicate that the heating time of a 0.1-μm-diameter aluminum particle may be of order 0.4 μs, whereas overall kinetics takes 10 μs: thus, the Al/water combination looks practical in principle for microrocket chambers characterized by short residence times.
124 citations
TL;DR: In this paper, a review of aero-gas-turbine engine high-pressure turbine performance degradation and the mechanisms that promote these losses are presented, along with specifications for next-generation engine clearance control systems.
Abstract: Improved blade-tip sealing in a high-pressure compressor and high-pressure turbine can provide dramatic improvements in specific fuel consumption, time on wing, compressor stall margin, and engine efficiency as well as increased payload and mission range capabilities. Maintenance costs to overhaul large commercial gas turbine engines can easily exceed $1 million. Removal of engines from service is primarily due to the spent exhaust gas temperature margin caused mainly by the deterioration of high-pressure-turbine components. Increased blade-tip clearance is a major factor in hot-section component degradation. As engine designs continue to push the performance envelope with fewer parts and the market drives manufacturers to increase service life, the need for advanced sealing continues to grow. A review of aero-gas-turbine engine high-pressure-turbine performance degradation and the mechanisms that promote these losses are presented. Benefits to the high-pressure turbine due to improved clearance management are identified. Past and present sealing technologies are presented along with specifications for next-generation engine clearance control systems.
116 citations
TL;DR: The In-FEEP thruster is a micro-propulsion device for the 1-100 μN thrust range with low thrust noise and high resolution as discussed by the authors, which is very important for many upcoming missions that require ultraprecise drag-free control such as the Gravity Field and Steady State Ocean Circulation Mission, LISA, Terrestrial Planet Finder/Darwin, or SMART-2.
Abstract: Indium liquid metal ion sources have been flying for more than 10 years on a variety of spacecraft for spacecraft potential control and as the core element of mass spectrometers. Since 1995, a dedicated indium field emission electric propulsion (In-FEEP) thruster has been under development and recently passed a 2000-h endurance test. The In-FEEP thruster is a micropropulsion device for the 1-100 μN thrust range with low thrust noise and high resolution. The latest performance characteristics including direct thrust measurements and beam profiles are summarized. This information is very important for many upcoming missions that require ultraprecise drag-free control such as the Gravity Field and Steady-State Ocean Circulation Mission, LISA, Terrestrial Planet Finder/Darwin, or SMART-2.
108 citations
TL;DR: In this article, wind-tunnel testing of a hypersonic inlet with a rectangular-to-elliptical shape transition has been conducted at Mach 4.0.
Abstract: Wind-tunnel testing of a hypersonic inlet with rectangular-to-elliptical shape transition has been conducted at Mach 4.0. This fixed geometry inlet had a geometric contraction ratio of 4.8 and was designed using a quasi-streamline tracing technique to have a design point of Mach 5.7. These tests were performed to investigate the starting and backpressure limits of the inlet at conditions well below its design point. Results showed that the inlet required side spillage holes in order to self-start at Mach 4.0. Once started, the inlet generated a compression ratio of 12.6, captured almost 80% of available air and withstood a backpressure ratio of 30.3 relative to tunnel static pressure. The spillage penalty for self-starting was estimated to be 3.4% of available air. These experimental results, along with previous experimental results at Mach 6.2, indicate that fixed-geometry inlets with rectangular-to-elliptical shape transition are a viable configuration for airframe-integrated scramjets that operate over a significant Mach-number range.
TL;DR: The development of a high-fidelity aerodynamic design optimization tool based on evolutionary algorithms for turbomachinery is attempted and a three-dimensional Navier-Stokes solver was used for aerodynamic analysis, showing the superiority of the present method over the conventional design approach.
Abstract: The development of a high-fidelity aerodynamic design optimization tool based on evolutionary algorithms for turbomachinery is attempted. A three-dimensional Navier-Stokes solver was used for aerodynamic analysis, so thatflowfields would be represented accurately and so that realistic and reliable designs would be produced. For efficient and robust design optimization, the real-coded adaptive range genetic algorithm was adopted, and the computation was parallelized and performed on an SGI Origin 2000 cluster to reduce turnaround time. The aerodynamic redesign of the NASA rotor 67 blade demonstrated the superiority of the present method over the conventional design approach, increasing adiabatic efficiency by 2% over the original design. This increase is achieved not only at the design condition, but over the entire operating range. This design optimization method has proven to be suitable for parallel computing. This promising tool is shown to help turbomachinery designers to design higher-performance machines while shortening the design cycle and reducing design costs.
TL;DR: In this paper, the authors present the design and manufacturing process innovations responsible for the remarkable evolution of the gasturbine engine, which has been credited with making the world much smaller, as well as with being a great benefitto humankind.
Abstract: Many of the design and manufacturing process innovations responsible for the remarkable evolution of the gasturbine engine, which has been credited with making the world much smaller, as well as with being a great benefitto humankind are presented. Many of the technologies presented show, from a design engineer's perspective, the levels of insight and foresight put forth by government agencies, the engineering community, and their leadership. The information presented offers a detailed understanding of technical issues that are not well known even within the industry. Universities can use this paper to bring the art of design into the classroom for applications of thermodynamics, materials, and structural design.
TL;DR: In this paper, three potential origins of side loads were observed and investigated, namely, the pressure fluctuations in the separation and recirculation zone due to the unsteadiness of the separation location, the transition of separation pattern between free-shock separation and restricted shock separation, and aeroelastic coupling, which indeed cannot cause but do amply existing side loads to significant levels.
Abstract: The operation of rocket engines in the overexpanded mode, that is, with the ambient pressure considerably higher than the nozzle exit wall pressure, can result in dangerous lateral loads acting on the nozzle. These loads occur as the boundary layer separates from the nozzle wall and the pressure distribution deviates from its usual axisymmetric shape. Different aerodynamic or even coupled aerodynamic/structural mechanic reasons can cause an asymmetric pressure distribution. A number of subscale tests have been performed, and three potential origins of side loads were observed and investigated, namely, the pressure fluctuations in the separation and recirculation zone due to the unsteadiness of the separation location, the transition of separation pattern between free-shock separation and restricted-shock separation, and aeroelastic coupling, which indeed cannot cause but do amply existing side loads to significant levels. All three mechanisms are described in detail, and methods are presented to calculate their magnitude and pressure ratio at which they occur.
TL;DR: The use of 3,6-bis(1H-1,2,3,4,4-tetrazol-5-ylamino)-s-Tetrazine (BTATz) and mixed Noxides (3,3'-azo-bis (6-amino)-1, 2, 4, 5-yamino-1.5) (DAATO3.5, where the 3.5 indicates the average oxide content) as solid fuels in micropropulsion systems is investigated.
Abstract: The use of two novel materials, 3,6-bis(1H-1,2,3,4-tetrazol-5-ylamino)-s-tetrazine (BTATz) and mixed N-oxides of 3,3'-azo-bis(6-amino-1,2,4,5-tetrazine) (DAATO3.5, where the 3.5 indicates the average oxide content), as solid fuels in micropropulsion systems is investigated. These materials were selected due to their ease of ignition, relatively good safety characteristics, long-term storage capability, low combustion temperature, noncorrosive combustion products, and reasonable projected material cost. Material safety data including impact, friction, differential scanning calorimetry, and electrostatic sensitivity are reported and compared to the previously implemented micropropulsion fuels, lead styphnate, glycidyl azide polymer (GAP) and 3,6-diamino-1,2,4,5-tetrazine-1,4-dioxide. Burn rate data are also reported for both materials. Whereas BTATz and DAATO3.5 have very high burn rates, DAATO3.5 is believed to have the highest burn rate of any known stable organic solid. Additionally, DAATO3.5 has a very desirable pressure exponent of 0.28. Test stand results using a quartz microthruster are reported for these materials. Estimated chamber pressures reached 18 atm with peak thrust levels around 0.1 N. Measured specific impulses of approximately 14% of the theoretical specific impulse (218 s for BTATz and 228 s for DAATO3.5) were obtained. This level of inefficiency is believed to be acceptable for micropropulsion systems and is comparable to other, independent work.
TL;DR: In this article, a series of detonation experiments conducted to characterize the deflagration-to-detonation transition (DDT) process for ethylene-air mixtures in a 44mm-square, 1.65m-long tube are described.
Abstract: The results from a series of detonation experiments conducted to characterize the deflagration-to-detonation transition (DDT) process for ethylene-air mixtures in a 44-mm-square, 1.65-m-long tube are described. Experiments were conducted for both single-shot detonations involving quiescent mixtures as well as multicycle detonations involving dynamic fill. For the experiments, high-frequency pressure and flame emission measurements were made to obtain the compression wave and flame speeds, respectively. In addition, schlieren and hydroxyl-radical/planar-laser-induced-fluorescence (OH-PLIF) imaging were applied to investigate the interactions between the shock-wave and combustion phenomena during both deflagration and detonation. For ethylene-air mixtures, strategically placed obstacles were necessary to achieve DDT. The effect of the presence of obstacles on flame acceleration was systematically investigated by changing the obstacle configuration. The parametric study of obstacle blockage ratio, spacing between obstacles, and length of the obstacle configuration indicated that for successful detonations the obstacle needs to accelerate the flame to a minimum flame speed of roughly half the Chapman-Jouguet detonation velocity. Differences in the flame and compression wave velocities demonstrated the development of a coupled feedback mechanism as the wave propagated along the tube. A series of simultaneous schlieren and OH-PLIF images showed that the obstacle plays a major role in generating small/large-scale turbulence that enhances flame acceleration. Localized explosions of pockets of unburned mixture further enhanced the shock-wave strength to continuously increase the flame speed. The results of this experimental study support the importance of obstacles as a means to enhance DDT and provide a potential solution for practical pulse-detonation-engine applications.
TL;DR: In this paper, exothermic laser ablation "fuels" for the micro Laser Plasma Thruster (µLPT), a novel type of microthruster, were investigated.
Abstract: We investigated exothermic laser ablation “fuels” for the micro Laser Plasma Thruster (µLPT), a novel type of microthruster. Using ms-duration laser pulses, which are required for multi-mode laser diodes to exceed ablation threshold fluence in the smallest focal spots available with conventional optics, successful target materials were restricted to those of low thermal conductivity, i.e., polymers, not metals. Polymers studied included carbon-doped polyvinylchloride for a passive target baseline, and several carbon-doped exothermic photopolymers specifically designed for their task, including polyvinylalcohol, a triazene polymer and a proprietary exothermic polymer (EP). In our single-shot impulse test setup, millimetric fuel samples were evaluated using a tiny torsion pendulum. Promising polymers were then made into fuel tapes and tested for continuous thrust under repetitive-pulse excitation. Two-layer fuel tapes consisted of a transparent supporting layer through which the light passed to ignite an absorbing fuel layer which formed a jet on the opposite side of the tape from that illuminated by the laser, an example of confined ablation. Best results were obtained with EP-1 up to 680 µN thrust with 2.1 W average optical power incident and jet velocity of 2‐3 km/s. Repeatability of our thrust-measuring torsion pendulum was improved to 1 µN.
TL;DR: A comparison of the combustion characteristics of aluminum hydride and aluminum with respect to combustion time and temperature at elevated pressures in carbon dioxide and oxygen has been performed using a shock tube.
Abstract: A comparison of the combustion characteristics of aluminum hydride and aluminum with respect to combustion time and temperature at elevated pressures in carbon dioxide and oxygen has been performed using a shock tube. Fo ra wide range of oxidizer concentrations, the combustion time of 5‐10 µ aluminum hydride is very similar to that of aluminum. Gas-phase temperatures measured by aluminum monoxide (AlO) emission and absorption spectroscopy, as well as solid product temperatures measured by two-color pyrometry, yield very similar values for both materials (to within experimental uncertainty). Observations of hydroxyl (OH) emission behind the incident shock suggest that hydrogen desorbs at temperatures well below the ignition threshold for aluminum. These observations are all consistent with the oxidation of aluminum hydride involving a rapid dehydrogenation step, followed by combustion of the remaining aluminum.
TL;DR: A detailed kinetic model devoted to the gas phase hypergolic ignition of MMH/NTO mixtures below 298 K and their combustion above 1000 K is presented in this article.
Abstract: A detailed kinetic model devoted to the gas phase hypergolic ignition of MMH/NTO mixtures below 298 K and their combustion above 1000 K is presented in this study. It consists of 403 equilibrated reactions among 82 species. This mechanism has been confronted with theoretical data available in the literature. The agreement between theory and predictions is found to be good. Important reactions for ignition at low initial temperature and pressure have been identified through sensitivity analyses. Two competing pathways can explain hypergolic ignition at low temperature. The formation of molecular preignition products is shown to promote the ignition. This conclusion was unexpected, as the formation of these products was generally considered to inhibit the ignition, by reactant consumption. Important reactions for combustion above 1000 K have been preliminarily identified. Some reactions are important both for low-temperature ignition and combustion, whereas some others are important either for ignition only or combustion only. The study is focused on the need for three reduced kinetic models.
TL;DR: In this paper, a potential power system concept includes dual 100 kWe Brayton converters, a deployable pumped loop heat rejection subsystem, and a 400 Vac Power Management and Distribution (PMAD) bus.
Abstract: The Jupiter Icy Moons Orbiter (JIMO) mission is currently under study by the Office of Space Science under the Project Prometheus Program. JIMO is examining the use of Nuclear Electric Propulsion (NEP) to carry scientific payloads to three Jovian moons. A potential power system concept includes dual 100 kWe Brayton converters, a deployable pumped loop heat rejection subsystem, and a 400 Vac Power Management and Distribution (PMAD) bus. Many trades were performed in aniving at this candidate power system concept. System-level studies examined design and off-design operating modes, determined startup requirements, evaluated subsystem redundancy options, and quantified the mass and radiator area of reactor power systems from 20 to 200 kWe. In the Brayton converter subsystem, studies were performed to investigate converter packaging options, and assess the induced torque effects on spacecraft dynamics due to rotating machinery. In the heat rejection subsystem, design trades were conducted on heat transport approaches, material and fluid options, and deployed radiator geometries. In the PMAD subsystem, the overall electrical architecture was defined and trade studies examined distribution approaches, voltage levels, and cabling options.
TL;DR: In this paper, the plume density, electron temperature, and plasma potential data collected with a combination of triple langmuir probes and floating emissive probes in a low-power, four-engine Hall thruster cluster are presented.
Abstract: The Hall thruster cluster is an attractive propulsion approach for spacecraft requiring very high-power electric propulsion systems. Plasma density, electron temperature, and plasma potential data collected with a combination of triple langmuir probes and floating emissive probes in the plume of a low-power, four-engine Hall thruster cluster are presented. Simple analytical formulas are introduced that allow these quantities to be predicted downstream of a cluster based solely on the known plume properties of a single thruster. Nomenclature A = area of one electrode AS = surface area of sheath surrounding an electrode B = magnetic field strength E = electric field strength e = electron charge kb = Boltzmann’s constant me = electron mass mi = ion mass n = electron number density n0 = reference density Te = electron temperature Te,0 = reference electron temperature Vd2 =v oltage measured between triple probe electrodes 1 and 2 Vd3 =v oltage applied between triple probe electrodes 1 and 3 V f = floating potential γ = ratio of specific heats δ = sheath thickness λD = electron Debye length φ = plasma potential φT = thermalized potential Subscript j = contribution from an individual thruster
TL;DR: A detailed experimental study of separated nozzle flows has been conducted at the NASA Langley Research Center 16-Foot Transonic Tunnel Complex as discussed by the authors, where force, moment, and pressure measurements were made and schlieren flow visualization was obtained for a subscale, nonaxisymmetric, two-dimensional, convergent-divergent nozzle.
Abstract: A detailed experimental study of separated nozzle flows has been conducted at the NASA Langley Research Center 16-Foot Transonic Tunnel Complex. As part of a comprehensive static performance investigation, force, moment, and pressure measurements were made and schlieren flow visualization was obtained for a subscale, nonaxisymmetric, two-dimensional, convergent-divergent nozzle. For reference, experimental results were compared with theoretical predictions based on one-dimensional gasdynamics and an approximate integral momentum boundary-layer method. Results from this study indicate that overexpanded nozzle flow was dominated by shock-induced boundary-layer separation, which was divided into two distinct flow regimes: three-dimensional separation with partial reattachment and fully detached two-dimensional separation. The test nozzle was observed to go through a marked transition in passing from one regime to the other. In all cases, separation provided a significant increase in static thrust efficiency compared to the ideal prediction. Results suggest that, with controlled separation, the entire overexpanded range of nozzle performance would be within 16% of the peak thrust efficiency. By offering savings in weight and complexity over a conventional mechanical variable geometry exhaust system, a fixed geometry nozzle may be able to cover an entire flight envelope in some applications.
TL;DR: In this article, self-pressurization of cryogenic storage tanks due to heat leak through the thermal protection system is examined along with the performance of various pressure control technologies for application in microgravity environments.
Abstract: Self-pressurization of cryogenic storage tanks due to heat leak through the thermal protection system is examined along with the performance of various pressure control technologies for application in microgravity environments. Methods of pressure control such as fluid mixing, passive thermodynamic venting, and active thermodynamic venting are analyzed using the homogeneous thermodynamic model. Simplified equations suggested may be used to characterize the performance of various pressure control systems and to design space experiments.
TL;DR: In this article, a premixed prevaporized 150-kW model scale combustor is investigated with two sets of swirling blades that induce flow rotation in the same direction (corotative) or in the opposite direction (counterrotative).
Abstract: Instabilities in a premixed prevaporized 150-kW model scale combustor are investigated experimentally. The injector fed with liquid heptane and preheated air features two sets of swirling blades that induce flow rotation in the same direction (corotative) or in the opposite direction (counter-rotative). The flame is stabilized with swirl behind a dump. Instabilities occur in the low-frequency range around 400 Hz corresponding to a quarter-wave mode acoustic coupling of the system. Simultaneous measurements of pressure and heat-release oscillations and phase-locked CH chemiluminescence images are used to characterize the combustion dynamics. In both corotative (COS) and counter-rotative (CNS) cases, the reaction region moves closer to the injector when the flame becomes unstable by about one-third of the stabilization distance under normal operation. Experiments indicate that the two swirl configurations have distinct domains of instability. The instability boundary separating stable and unstable regions can be defined in terms of a critical velocity v c , which depends on the equivalence ratio Φ, air injection temperature T i n j , and swirl geometry. In the coswirl configuration, instabilities occur when the injection velocity is lower than the critical velocity [u v c (Φ, T i n j ; CNS)]. In a range of conditions corresponding to low injection velocities, reduced eqivalence ratio, and for the coswirl configuration, an unsteady flashback takes place in which the flame moves periodically in and out of the fuel premixer. This mechanism is related to the existence of a low-velocity region near the injector exit plane. Observations of the space-time development of the heat release under unstable operation indicate that the oscillations are significantly influenced by the swirl geometry and are caused by different mechanisms. The coswirl configuration features a central recirculation, which gives rise to periodic vortex roll-up, convection, and sudden release of heat when the vortices impinge on the lateral walls. In the counterswirl geometry there are no identifiable flow structure, but the heat-release pattern is convected periodically in the chamber. Estimates of the delay times associated with the two mechanisms support the view that coswirl instabilities are driven by vortex roll-up, whereas counterswirl instabilities are probably sustained by equivalence-ratio inhomogeneities.
TL;DR: In this paper, the derivation of a compressor characteristic, and the experimental validation of a dynamic model for a variable speed centrifugal compressor using this characteristic, are presented, where the dynamic compressor model of Fink et al. is used, and a variable-speed compressor characteristic is derived by the use of energy transfer and loss analysis.
Abstract: The derivation of a compressor characteristic, and the experimental validation of a dynamic model for a variable speed centrifugal compressor using this characteristic, are presented. The dynamic compressor model of Fink et al. is used, and a variable speed compressor characteristic is derived by the use of energy transfer and loss analysis. It is demonstrated that taking into account the losses due to friction, incidence, mixing, and blade loading results in compressor characteristics that closely match the measured characteristics. The simulated response of the dynamic model was found to be in excellent agreement with the experimental results, both for set point changes using fuel flow and blow off and for surge oscillations. Analysis of the power spectrum of the in-surge rotational speed and pressure oscillations reveal that the simulated nonlinear oscillations match experimental values up to the third harmonic, both with respect to frequency and amplitude.
TL;DR: In this paper, a finite element analysis model was developed to investigate the dynamic characteristics of single-and dual-rotor bearing turbomachinery systems, and the model was applied to both single and dual rotor bearing applications.
Abstract: Finite element analysis models were developed to investigate the dynamic characteristics of single- and dual-rotor-bearing turbomachinery systems. When an inertial coordinate system was used, the dynamic models of the rotor-bearing systems included gyroscopic moments, rotary inertias, and bending and shear deformations. The models were analyzed to predict the natural frequencies, to produce critical speed maps, and to estimate the bearing stiffnesses. These rotor-bearing system analyses were then applied to both single-rotor and dual-rotor system applications. In the single-rotor system application, a small turbojet engine and its rotor components were used as a basis for the model. Both theoretical and experimental analyses were used to study this engine rotor-bearing system. Modal testing and a dynamic engine test were used to verify the analytical results, including the predicted critical speed map and the bearing stiffnesses. Very good agreement was found between the analyses and the test data. In the dual-rotor application, the effects of the speed ratio of the high-speed to low-speed shafts of the dual-rotor system on the critical speeds was studied. It was demonstrated that this speed ratio could be used as one of the dual-rotor system design parameters. Finally, it was noted that the interrotor bearing stiffness between the high-speed and the low-speed shafts of the dual-rotor system affected the mode shapes of the shafts within the system, in addition to the rotor system critical speeds.
TL;DR: In this article, a 2-10 wt% MnO 2 supported on TiO 2 was examined for H 2 O 2 decomposition and a reaction mechanism involving an Mn 4 + -Mn 3 + redox couple has been proposed.
Abstract: Catalysts containing 2-10 wt% MnO 2 supported on TiO 2 were examined for H 2 O 2 decomposition. Catalysts were prepared by impregnation of a high TiO 2 surface area. X-ray defraction patterns, scanning electron micrographs, and surface area measurements gave evidence of a uniform distribution of the active phase on the support surface and of the absence of segregated manganese oxides phases. Temperature programmed reduction measurements showed the presence, beside MnO 2 , of different Mn oxides species formed by interaction between the active phase and the support surface. The MnO 2 reducibility increased, whereas the mean oxidation state of manganese decreased with increasing manganese content. Catalytic tests were performed in a batch reactor with 50 or 70% concentration H 2 O 2 solutions. Catalytic activity was very high at the beginning of the tests and decreased with time, reaching a final constant value that increased with manganese content. Kinetic constants, evaluated assuming a first-order reaction rate, were comparable or higher than those of similar manganese-based catalysts. An optimal MnO 2 content was found, corresponding to the quite complete surface monolayer coverage. A reaction mechanism involving an Mn 4 + -Mn 3 + redox couple has been proposed.
TL;DR: In this article, selected gel propellants and simulants were formulated, prepared, rheologically characterized, and tested in the first phase of a program to develop gel-propulsion technology infrastructure.
Abstract: Selected gel propellants and simulants were formulated, prepared, rheologically characterized, and tested in the first phase of a program to develop gel-propulsion technology infrastructure. Hydrazine-based fuels, gelled with polysaccharides, were characterized as shear-thinning pseudoplastic fluids with low-shear yield stress (τ y i e l d ), whereas inhibited red-fuming nitric acid (IRFNA) and hydrogen peroxide oxidizers, gelled with silica, were characterized as yield thixotropic fluids with significant τ y i e l d . Safe storage and handling procedures were established. A laboratory-scale experimental setup was used to hot fire successfully a small 100-N nominal thrust rocket engine with selected hypergolic neat-liquid and gelled bipropellant combinations. One-element pentad-type injectors were utilized in the tests to inject the propellants into the combustion chamber. Continuous tests of up to 25-s firing duration and multipulse operations of up to 20 cycles of 0.1-s on/0.5-s off were successfully conducted with gelled-hydrazine/IRFNA bipropellants. Neat-liquid and gelled mono-methyl hydrazine/IRFNA bipropellants were also tested. The combustion pressure ranged between 20 and 35 bars. Experimental characteristic velocity, c* e x p , was determined as a function of the oxidizer-to-fuel (O/F) mass flow rate ratio. Maximum c* efficiency of more than 95 and about 90% was obtained in continuous firings for the neat-liquid and gelled hydrazine/IRFNA, respectively. In both cases, the maximum c* e x p values were obtained at higher O/F ratios than those that yield maximum theoretical c*.
TL;DR: In this article, a wavelet transform is used to analyze the dynamic pressure signals of a low-speed axial compressor, and it is found that the emergence of spikes is strongly related to two structural asymmetries of this compressor, one in the rotating frame and the other in the stationary frame.
Abstract: Dynamic pressure signals taken from a low-speed axial compressor are analyzed with a wavelet transform. Several practical issues of continuous wavelet transforms, within the context of rotor-tip flow analysis, are discussed. The discrete form of continuous wavelet transform is presented with a graphical derivation that uniquely explains a complicated mathematical process with a readable graphical language without losing much mathematical strictness. With this wavelet tool, the data from a low-speed, three-stage axial compressor are analyzed. In these data, the wavelet analysis is able to pinpoint the exact time when spikes were initiated, as early as hundreds of rotor re volutions prior to stall, and track them thereafter. It is found that the emergence of the spikes is strongly related to two structural asymmetries of this compressor, one in the rotating frame and the other in the stationary frame. The spike development is not continuous. The ability to track the prestall process in both time and space, as demonstrated in this paper, makes wavelet analysis an effective tool for investigating the complicated instability phenomena of axial compressors.
TL;DR: In this article, the effectiveness of an H 2 /O 2 torch and a plasma torch in promoting the ignition of a supersonic combustor was investigated analytically and experimentally.
Abstract: The effectiveness of an H 2 /O 2 torch igniter, which injects H 2 /O 2 combustion gas, and a plasma torch igniter in promoting ignition of an H 2 /air mixture within a supersonic combustor was investigated analytically and experimentally. First, bulk temperature and bulk radical concentration in an O 2 plasma torch igniter were estimated by asymptotic analysis. Those in the H 2 /O 2 torch igniter were also obtained by equilibrium calculation. Second, gas sampling was conducted to determine the concentration of igniter injectant and the equivalence ratio in the region of ignition, namely, the recirculation region at the base of a rearward-facing step. Then, ignition time of the mixture in the region was analytically estimated by using a two-step H 2 /O 2 reaction mechanism and the results just-mentioned. The estimation showed that the ignition promotion effect of the plasma torch igniter was higher than that of the H 2 /O 2 torch igniter at a given input energy. However, the H 2 /O 2 torch igniter was also expected to sufficiently promote ignition because it was easy to increase the input energy by increasing the mass flow rate. Thus, an ignition experiment with the H 2 /O 2 torch igniter was conducted in Mach 2.5 airflow. Comparison of the results with those of the plasma torch igniters showed that the H 2 /O 2 torch igniter also had a sufficient ignition promotion effect as expected from analytical results.