Showing papers on "Rocket published in 2012"

Design of liquid-propellant rocket engines

Abstract: A design method for liquid-propellant rocket engines is proposed, on the basis of models adjusted in response to test results. This method increases the reliability of the results and reduces ground testing of the motors. The effectiveness of the method is confirmed by ground testing and flight experience with the motors developed by Salyut Design Bureau at Krunichev State Space-Research Center.

Ground and Space-Based Measurement of Rocket Engine Burns in the Ionosphere

Abstract: On-orbit firings of both liquid and solid rocket motors provide localized disturbances to the plasma in the upper atmosphere. Large amounts of energy are deposited to ionosphere in the form of expanding exhaust vapors which change the composition and flow velocity. Charge exchange between the neutral exhaust molecules and the background ions (mainly O+) yields energetic ion beams. The rapidly moving pickup ions excite plasma instabilities and yield optical emissions after dissociative recombination with ambient electrons. Line-of-sight techniques for remote measurements rocket burn effects include direct observation of plume optical emissions with ground and satellite cameras, and plume scatter with UHF and higher frequency radars. Long range detection with HF radars is possible if the burns occur in the dense part of the ionosphere. The exhaust vapors initiate plasma turbulence in the ionosphere that can scatter HF radar waves launched from ground transmitters. Solid rocket motors provide particulates that become charged in the ionosphere and may excite dusty plasma instabilities. Hypersonic exhaust flow impacting the ionospheric plasma launches a low-frequency, electromagnetic pulse that is detectable using satellites with electric field booms. If the exhaust cloud itself passes over a satellite, in situ detectors measure increased ion-acoustic wave turbulence, enhanced neutral and plasma densities, elevated ion temperatures, and magnetic field perturbations. All of these techniques can be used for long range observations of plumes in the ionosphere. To demonstrate such long range measurements, several experiments were conducted by the Naval Research Laboratory including the Charged Aerosol Release Experiment, the Shuttle Ionospheric Modification with Pulsed Localized Exhaust experiments, and the Shuttle Exhaust Ionospheric Turbulence Experiments.

CFD analysis of transcritical methane in rocket engine cooling channels

Abstract: The knowledge of the flow behavior inside cooling channels is of great importance to improve design and performance of regeneratively cooled rocket engines. The modeling of the coolant flow is a challenging task because of its particular features, such as the high wall temperature gradient, the high Reynolds number and the three-dimensional geometry of the passages. In case of methane as coolant, a further complication is the transcritical operating condition of the fluid. In this thermodynamic regime large changes of the fluid properties can greatly influence the coolant flow-field and the heat transfer. Numerical simulations of transcritical methane flow-field in asymmetrically heated rectangular channel with high aspect ratio and strong wall temperature differences are carried out by a suitable CFD solver. Results are discussed in detail and compared with supercritical methane flow-fields. Finally the aspect ratio effect on methane flow is analyzed by comparison of four different rectangular cooling channel geometries with fixed hydraulic diameter and coolant flows with the same mass flow-rate per unit area. Emphasis is given to the comparison of fluid cooling performance and pressure loss.

Neutral Polymeric Bonding Agents (NPBA) and Their Use in Smokeless Composite Rocket Propellants Based on HMX-GAP-BuNENA

Abstract: Landsem, Eva; Jensen, Tomas Lunde; Hansen, Finn Knut; Unneberg, Erik; Kristensen, Tor Erik Holt. Neutral Polymeric Bonding Agents (NPBA) and Their Use in Smokeless Composite Rocket Propellants Based on HMXGAP- BuNENA. Propellants, explosives, pyrotechnics 2012 ;Volum 37.(5) s. 581-591

Nuclear Thermal Propulsion (NTP): A proven growth technology for human NEO/Mars exploration missions

Abstract: The nuclear thermal rocket (NTR) represents the next “evolutionary step” in high performance rocket propulsion. Unlike conventional chemical rockets that produce their energy through combustion, the NTR derives its energy from fission of Uranium-235 atoms contained within fuel elements that comprise the engine's reactor core. Using an “expander” cycle for turbopump drive power, hydrogen propellant is raised to a high pressure and pumped through coolant channels in the fuel elements where it is superheated then expanded out a supersonic nozzle to generate high thrust. By using hydrogen for both the reactor coolant and propellant, the NTR can achieve specific impulse (I sp ) values of ∼900 seconds (s) or more — twice that of today's best chemical rockets. From 1955–1972, twenty rocket reactors were designed, built and ground tested in the Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) programs. These programs demonstrated: (1) high temperature carbide-based nuclear fuels; (2) a wide range of thrust levels; (3) sustained engine operation; (4) accumulated lifetime at full power; and (5) restart capability — all the requirements needed for a human Mars mission. Ceramic metal “cermet” fuel was pursued as well, as a backup option. The NTR also has significant “evolution and growth” capability. Configured as a “bimodal” system, it can generate its own electrical power to support spacecraft operational needs. Adding an oxygen “afterburner” nozzle introduces a variable thrust and I sp capability and allows bipropellant operation. In NASA's recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, versatile vehicle design, simple assembly, and growth potential. In contrast to other advanced propulsion options, no large technology scale-ups are required for NTP either. In fact, the smallest engine tested during the Rover program — the 25,000 lb f (25 klb f ) “Pewee” engine is sufficient when used in a clustered engine arrangement. The “Copernicus” crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth object (NEO) and Mars orbital missions prior to a Mars landing mission. The paper also discusses NASA's current activities and future plans for NTP development that include system-level Technology Demonstrations — specifically ground testing a small, scalable NTR by 2020, with a flight test shortly thereafter.

Visualization of the Liquid Layer Combustion of Paraffin Fuel for Hybrid Rocket Applications

Abstract: Cantwell, B., Karabeyoglu, A., & Altman, D. Recent Advances in Hybrid Propulsion. Int. J of Energetic Mat and Chem Prop. Vol. 9, Art. 1344, 2010. Haapanen, Siina Ilona. Linear Stability Analysis and Direct Numerical Simulation of a Miscible Two-Fluid Channel Flow. Diss. Stanford University, Stanford, CA, May 2008. Bibliography Researchers at Stanford University have identified a class of paraffin-based fuels that burn at regression rates that are 3-4 times that of conventional hybrid fuels. These fuels produce a thin liquid surface layer when they burn. The liquid layer has low surface tension and viscosity and thus is unstable under the oxidizer gas flow, leading to entrainment of droplets into the gas stream. The entrainment mechanism effectively acts like a spray injection system distributed along the fuel port, greatly increasing the mass transfer rate of the fuel and hence the fuel regression rate. The total fuel regression rate is comprised of contributions from the entrainment mass transfer and from evaporation mass transfer. The entrainment component has some dependence on the chamber pressure in addition to being a function of the surface tension, thickness and viscosity of the melt layer. Not all fuels that form a melt layer at their surface will entrain. Low surface tension and low viscosity within the melt layer are the key criterion for entrainment and hence for identifying high regression rate fuels. Liquid Layer Combustion Cantwell, Karabeyoglu, & Altman, 2009 Diffusion Flame Entrained Droplets

Wall Pressure Unsteadiness and Side Loads in Overexpanded Rocket Nozzles

Abstract: Surveys of both the static and dynamic wall pressure signatures on the interior surface of a sub-scale, cold-flow and thrust optimized parabolic nozzle are conducted during fixed nozzle pressure ratios corresponding to FSS and RSS states. The motive is to develop a better understanding for the sources of off-axis loads during the transient start-up of overexpanded rocket nozzles. During FSS state, pressure spectra reveal frequency content resembling SWTBLI. Presumably, when the internal flow is in RSS state, separation bubbles are trapped by shocks and expansion waves; interactions between the separated flow regions and the waves produce asymmetric pressure distributions. An analysis of the azimuthal modes reveals how the breathing mode encompasses most of the resolved energy and that the side load inducing mode is coherent with the response moment measured by strain gauges mounted upstream of the nozzle on a flexible tube. Finally, the unsteady pressure is locally more energetic during RSS, albeit direct measurements of the response moments indicate higher side load activity when in FSS state. It is postulated that these discrepancies are attributed to cancellation effects between annular separation bubbles.

Rocket or ballistic launch rotary wing vehicle

Abstract: Embodiments of the present invention relate to a rocket or ballistic launch rotary wing air vehicle.

Nammo Hybrid Rocket Propulsion TRL Improvement Program

Abstract: In 2010, Nammo initiated an ESA-funded hybrid propulsion Technology Readiness Level (TRL) improvement program based on ideas proposed in a study performed for Ariane 5 ME a year before. In a program lasting less than one year, Nammo executed 29 hybrid rocket firing tests using hydrogen peroxide as the oxidizer. All the tests were successful, demonstrating hybrid rocket propulsion exhibiting high performance under different conditions.

Feasibility Study and Demonstration of an Aluminum and Ice Solid Propellant

Abstract: Aluminum-water reactions have been proposed and studied for several decades for underwater propulsion systems and applications requiring hydrogen generation. Aluminum and water have also been proposed as a frozen propellant, and there have been proposals for other refrigerated propellants that could be mixed, frozen in situ, and used as solid propellants. However, little work has been done to determine the feasibility of these concepts. With the recent availability of nanoscale aluminum, a simple binary formulation with water is now feasible. Nanosized aluminum has a lower ignition temperature than micron-sized aluminum particles, partly due to its high surface area, and burning times are much faster than micron aluminum. Frozen nanoscale aluminum and water mixtures are stable, as well as insensitive to electrostatic discharge, impact, and shock. Here we report a study of the feasibility of an nAl-ice propellant in small-scale rocket experiments. The focus here is not to develop an optimized propellant; however improved formulations are possible. Several static motor experiments have been conducted, including using a flight-weight casing. The flight weight casing was used in the first sounding rocket test of an aluminum-ice propellant, establishing a proof of concept for simple propellant mixtures making use of nanoscale particles.

Flight demonstration of a milliarcsecond pointing system for direct exoplanet imaging.

Abstract: We present flight results from the optical pointing control system onboard the Planetary Imaging Concept Testbed Using a Rocket Experiment (PICTURE) sounding rocket PICTURE (NASA mission number: 36225 UG) was launched on 8 October 2011, from White Sands Missile Range It attempted to directly image the exozodiacal dust disk of ϵ Eridani (K2V, 322 pc) down to an inner radius of 15 AU using a visible nulling coronagraph The rocket attitude control system (ACS) provided 627 milliarcsecond (mas) RMS body pointing (~2'' peak-to-valley) The PICTURE fine pointing system (FPS) successfully stabilized the telescope beam to 51 mas (002λ/D) RMS using an angle tracker camera and fast steering mirror This level of pointing stability is comparable to that of the Hubble Space Telescope We present the hardware design of the FPS, a description of the limiting noise sources and a power spectral density analysis of the FPS and rocket ACS in-flight performance

Peregrine Hybrid Rocket Motor Ground Test Results

Abstract: To further develop and demonstrate the applicability of liquefying-fuel hybrid rocket technology to low-cost launch applications, a small team of engineers is developing a medium-scale liquefying-fuel hybrid sounding rocket using storable propellants (paraffin wax and N2O) that will carry a 5 kg payload to the edge of space. This rocket, known as Peregrine, is being developed by engineers from NASA Ames, Stanford University, Space Propulsion Group Inc. (SPG, Sunnyvale, CA) and NASA Wallops, with a launch from Wallops anticipated at some point in the future. This paper focuses on the propulsion ground test results obtained to date.

PICTURE: a sounding rocket experiment for direct imaging of an extrasolar planetary environment

Abstract: The Planetary Imaging Concept Testbed Using a Rocket Experiment (PICTURE 36.225 UG) was designed to directly image the exozodiacal dust disk of ǫ Eridani (K2V, 3.22 pc) down to an inner radius of 1.5 AU. PICTURE carried four key enabling technologies on board a NASA sounding rocket at 4:25 MDT on October 8th, 2011: a 0.5 m light-weight primary mirror (4.5 kg), a visible nulling coronagraph (VNC) (600-750 nm), a 32x32 element MEMS deformable mirror and a milliarcsecond-class fine pointing system. Unfortunately, due to a telemetry failure, the PICTURE mission did not achieve scientific success. Nonetheless, this flight validated the flight-worthiness of the lightweight primary and the VNC. The fine pointing system, a key requirement for future planet-imaging missions, demonstrated 5.1 mas RMS in-flight pointing stability. We describe the experiment, its subsystems and flight results. We outline the challenges we faced in developing this complex payload and our technical approaches.

Experimental investigation of high frequency combustion instability in cryogenic oxygen-hydrogen rocket engines.

Abstract: Self-sustaining pressure oscillations in the combustion chamber, or combustion instability, is a commonly encountered and potentially damaging phenomenon in liquid propellant rocket engines (LPREs). In the high-frequency variety of combustion instability, the pressure oscillations take on the form and frequency of an acoustic resonance mode of the combustion chamber volume. The most common mode in naturally occurring instability, and also the most destructive, is the first tangential mode, with acoustic gas oscillations oriented transversally to the direction of propellant injection. The instability is driven by the coupling between acoustic oscillations and unsteady energy release from combustion. The mechanisms through which injection and combustion firstly respond to the acoustic field, and secondly feed energy back into the acoustic field have not yet been fully characterised.

Mechanical Properties of Smokeless Composite Rocket Propellants Based on Prilled Ammonium Dinitramide

Abstract: Landsem, Eva; Jensen, Tomas Lunde; Hansen, Finn Knut; Kristensen, Tor Erik Holt; Unneberg, Erik. Mechanical Properties of Smokeless Composite Rocket Propellants Based on Prilled Ammonium Dinitramide. Propellants, explosives, pyrotechnics 2012 ;Volum 37.(6) s. 691-698

Design and optimization of variable thrust hybrid rocket motors for sounding rockets

Abstract: Besides safety and low-cost, the start/shutdown/restarting and throttling ability are the other two significant advantages of hybrid rocket motors (HRMs) compared with liquid and solid ones. In this study, a two-stage variable thrust and non-toxic 98%HP/HTPB hybrid rocket motor (VTHRM) is designed and applied in a sounding rocket, and the design parameters of the motor are analyzed and optimized. A computational program is developed to design the motor system structure, to predict the interior ballistics and the ballistic trajectory. A star grain and a wheel grain are compared. The design of experiment (DOE), variance analysis and the main effect analysis are employed to investigate the influence of the main design parameters on motor performance. The multidiscipline feasible (MDF) approach is applied to establish the optimization procedure after analyzing the system design structure matrix. A modified differential evolution algorithm is employed to maximize the load mass. The results indicate that the wheel grain could obtain a larger load mass and a lower length to diameter ratio, and that throttling markedly meliorates the motor and rocket performance. The conclusions drawn from the analysis and optimization could provide instructive guide and theoretical basis for engineering designs.

Unsteady PSP Measurement of Transonic Unsteady Flow Field around a Rocket Fairing Model

Abstract: Unsteady Pressure-Sensitive Paint (PSP) measurement of a rocket faring model was conducted at the JAXA 2m×2m Transonic Wind Tunnel as a part of the campaign to investigate the steady and unsteady flow field and to obtain the CFD validation data. Unsteady PSP measurement results of M=0.8 and =0 and +4° were described in this paper. Anodized-aluminum PSP (AA-PSP) was coated on the model surface, and it was illuminated by a high-power blue laser diode. AA-PSP luminescence was measured by a high-speed camera, whose frame rate was 5k frame-per-sec. The high-speed camera images were reduced to the time-series Cp distribution images using in-situ method involving pressure transducer data. Several kinds of data analysis were applied to the time-series Cp images and discussed about steady pressure distribution, unsteady pressure variation, and the results of the spectral analysis of the flow field around a rocket faring configuration.

Developments in the field of automatic guidance and control of rockets

Abstract: Introduction T definitions are that "navigation" determines the vehicle's state for initial conditions as well as during flight; that "guidance" selects the maneuvering sequence to get from the instantaneous state to a required state; and that "control" executes the maneuvers called for by guidance. These definitions did not exist in the early development years. Navigation corrections were usually taken care of by flight time deviations, which were included in simple guidance functions. Fixed programs, such as flight-tilt programs and roll programs, to rotate the vehicle into its required flight plane have been considered control functions rather than guidance functions and have been exceptions to these definitions. The development of liquid fuel and oxidizer engines made it necessary to stabilize the rockets and to provide attitude control. This included a pitch tilt program, because these rockets with their low initial acceleration had to be launched in a vertical direction. In order to aim at a target the next requirements were to constrain the rocket to a predetermined flight plane and to provide a well-defined velocity vector at propulsion cutoff of the rocket engine. Related but less sophisticated requirements had existed for ship and airplane control. Inertial sensors, such as gimbal suspended gyroscopes and rate gyros, had been used for these systems. In 1912 Prof. Max Schuler had designed a threegyro, three-degree-of-freedom gyro compass for ship navigation, and two-degree-of-freedom gyros were used later as airplane orientation sensors. Schemes and components derived from these systems were applied to the first rocket attitude control systems. A somewhat more difficult task was to obtain guidance sensors, which did not exist yet. For many years inertial sensors were in competition with radio guidance devices, the latter being more promising in the early years to fulfill accuracy requirements. Accelerometers had to be used as inertial sensors for guidance. Velocity and displacement, as the information required, had to be derived; the first with normalized accuracy demands of at least 10~-10~ . The derived signals had to be obtained by the only available method, which was mechanical integration. In contrast to the inertial sensors, the radio sensors measured lateral displacements and forward (radial) velocities directly. Since the guidance systems had to be used for missiles an inertial system was preferable because it is self-contained and not subjected to external interference as radio guidance systems are. The radio guidance system became available and operational first. It yielded an accuracy in range that was about 10% better than the inertial system introduced afterward. Only a few test flights with a pure inertial system took place before the end of World War II. The cross-range impact errors were approximately twice as large as with the radio guidance system. Thus for some time it

Computational studies of the effects of oxidiser injector length on combustion instability

Abstract: Computational studies were conducted to investigate the effects of oxidizer injector length on combustion instability for a single-eleme nt, longitudinal rocket combustor. The configuration is based on companion experiments that were developed to provide validation data for combustion instability. The experiment use s gaseous methane as fuel and decomposed hydrogen peroxide as oxidizer. As verified with multiple diagnostics, a baseline case demonstrates similar instability behavior as t he associated experimental case, with instability occurring at similar acoustic modes. Th is baseline case is then compared against four other oxidizer injector lengths, at similar ox idizer injector lengths as the experiment. The general trend of the computations is similar to the experiments, where the intermediate oxidizer injector length demonstrates the highest l evel of instability. The mechanisms which cause instability are then investigated using the c omputational results, with possible explanations given for the experimental behavior.

Transpiration-Cooled Ceramic Thrust Chamber Applicability for High-Thrust Rocket Engines

Abstract: The development of ceramic rocket engine thrust chambers at the German Aerospace Center (DLR) currently concentrates on designs of self-sustaining, transpiration-cooled, fiber-reinforced ceramic rocket engine chamber structures. This paper discusses characteristic issues and potential benefits introduced by this technology. Achievable benefits are the reduction of weight and manufacturing cost, as well as an increased reliability and higher lifetime due to thermal cycle stability. This paper discusses the current status of DLR's ceramic thrust chamber technology and potential applications for high thrust engines. The test results of DLR's ceramic thrust chamber project KSK are used for a rough approximation of the performance of high thrust applications. Based on the KSK test results an extrapolation is performed. c*-efficiency and geometrical scaling effects are taken into consideration. Due to favorable scaling effects, high thrust applications will profit by all benefits of the discussed technology, while avoiding the most significant performance drawbacks.

Noniterative Instantaneous Impact Point Prediction Algorithm for Launch Operations

Abstract: T HE instantaneous impact point (IIP) in launch operations is defined as the touchdown point of a rocket with an assumption of an immediate end of the propelled flight [1–3]. The IIP is very important information to judge whether the flight is normal or not, and it should be continuously monitored during the whole flight so that proper actions be taken in an emergency to protect human life and property. Therefore, the real-time prediction and monitoring of the IIP is one of the key safety requirements. The IIP prediction problem seeks for the intersection point between the Earth’s surface and the free-flight trajectory of a rocket subject to gravity and other disturbing forces. The current position and velocity vectors are given as inputs to the problem. A numerical integration of the dynamicmodel including all force elementswith an initial position and velocity could yield a very accurate answer to the problem with relative ease. But when the real-time calculation is considered as a constraint, the numerical integration that requires a lot of computational resources is not a feasible option in the majority of launch operations. There has been some research on developing the IIP prediction algorithm that provides the real-time solution within acceptable accuracy with limited computational resources. To calculate the IIP for shortto midrange rockets, simplified models in the localvertical–local-horizontal frame are frequently used [3–6]. These assume that the trajectory of a rocket is parabolic subject to the constant vertical downward gravitational acceleration and often add correction terms to compensate for the modeling errors. Montenbruck and Markgraf [6] developed a linearized IIP calculation algorithm and implemented it for the launch operation of the Maxus-5 sounding rocket. Their algorithm was based on a simple parabolic flight model, and its prediction accuracy was improved by considering the effects of the Earth curvature, the Coriolis force, and the gravity field variation. The Keplerian algorithm predicts the IIP using the equations describing the motion of a rocket under the inverse-square law of gravity. Regan and Anandakrishnan [4] and Zarchan [7] developed a series of formulations that can be used as the base of theKeplerian IIP prediction, such as the expressions for the flight angle and the flight time. Currently, an iterativeKeplerian algorithm is frequently used to predict the IIP for long-range space vehicles [8,9]. The iterative Keplerian algorithm can calculate the true IIP solution for the Keplerian motion of a rocket using an iterative procedure (e.g., successive substitution) to find parameters such as Kepler’s f and g series expansions, the eccentric anomaly of the impact point, and the surface radius at impact. Considering that the IIP prediction for an actual launch operation should be carried out in real time, the iterative procedure that has the risk of computational overload is not very desirable to be included in the prediction algorithm. Especially when the algorithm is implemented in an onboard computer that has limited computing resources but is responsible for multiple mission-critical tasks such as event sequencing, navigation, and attitude control, the potential damage associated with the risk is significantly increased. This engineering Note presents a Keplerian IIP prediction algorithm that does not require any iterative procedure and is thus remarkably faster than the iterative Keplerian algorithm. First, an exact and closed solution for the Keplerian IIP with the spherical Earth model is developed. Then, a noniterative procedure to compensate for the effect of the oblateness of the Earth is added. Finally, the validity of the proposed algorithm is proved through the performance comparison with the iterative algorithm in terms of computational efficiency and the prediction accuracy.

Frequency-Modulated Pulse-Jet Control of an Artillery Rocket

Abstract: Longandmedium-range artillery rockets are used for indirect fire on distant targets. As they have a large impactpoint dispersion, they are considered tobe areaweapons.Nevertheless, a control system is required in order to reduce the dispersion and increase hit probability. The pulse-jet control system is simple and efficient enough to produce acceptable impact-point dispersion. However, it is necessary to perform an optimization of the control logic in order to obtain a satisfactory performance with the minimum number and intensity of control pulses. A new two-phase guidance scheme named trajectory tracking with pulse-frequencymodulation is presented, considering the fact that an artillery rocket flies through different atmospheric environments. Simulation studies have been conducted in order to perform parametric and performance analyses. It is shown that the presented guidance scheme achieves excellent accuracy even in the case of a small number of pulse jets. In the case of a large number, it tends to nullify the impact-point dispersion. It is also shown that in comparison with the window-based trajectory-tracking guidance, the proposed method achieves significantly better results.

Highlights From a Mach 4 Experimental Demonstration of Inlet Mode Transition for Turbine-Based Combined Cycle Hypersonic Propulsion

Abstract: NASA is focused on technologies for combined cycle, air-breathing propulsion systems to enable reusable launch systems for access to space. Turbine Based Combined Cycle (TBCC) propulsion systems offer specific impulse (Isp) improvements over rocket-based propulsion systems in the subsonic takeoff and return mission segments along with improved safety. Among the most critical TBCC enabling technologies are: 1) mode transition from the low speed propulsion system to the high speed propulsion system, 2) high Mach turbine engine development and 3) innovative turbine based combined cycle integration. To address these challenges, NASA initiated an experimental mode transition task including analytical methods to assess the state-of-the-art of propulsion system performance and design codes. One effort has been the Combined-Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE-LIMX) which is a fully integrated TBCC propulsion system with flowpath sizing consistent with previous NASA and DoD proposed Hypersonic experimental flight test plans. This experiment was tested in the NASA GRC 10 by 10-Foot Supersonic Wind Tunnel (SWT) Facility. The goal of this activity is to address key hypersonic combined-cycle engine issues including: (1) dual integrated inlet operability and performance issues-unstart constraints, distortion constraints, bleed requirements, and controls, (2) mode-transition sequence elements caused by switching between the turbine and the ramjet/scramjet flowpaths (imposed variable geometry requirements), and (3) turbine engine transients (and associated time scales) during transition. Testing of the initial inlet and dynamic characterization phases were completed and smooth mode transition was demonstrated. A database focused on a Mach 4 transition speed with limited off-design elements was developed and will serve to guide future TBCC system studies and to validate higher level analyses.

Hypergolic Systems: A Review in Patents

Abstract: Hypergolic reactions may be useful in civil and military applications. In the area of rocket propulsion, they constitute a potential oeld due to the reduced Zeight and comple(ity of fuel inMection systems, alloZing yet controlled use of the propulsors. This manuscript aimed at presenting different hypergolic systems and their particularities, comparing them Zith chemical propulsion systems, Zhich are most commonly employed in rocket motors, for e(ample.