Showing papers on "Rocket published in 2011"

Experimental research on the rotating detonation in gaseous fuels–oxygen mixtures

Abstract: An experimental study on rotating detonation is presented in this paper. The study was focused on the possibility of using rotating detonation in a rocket engine. The research was divided into two parts: the first part was devoted to obtaining the initiation of rotating detonation in fuel–oxygen mixture; the second was aimed at determination of the range of propagation stability as a function of chamber pressure, composition, and geometry. Additionally, thrust and specific impulse were determined in the latter stage. In the paper, only rich mixture is described, because using such a composition in rocket combustion chambers maximizes the specific impulse and thrust. In the experiments, two kinds of geometry were examined: cylindrical and cylindrical-conic, the latter can be simulated by a simple aerospike nozzle. Methane, ethane, and propane were used as fuel. The pressure–time courses in the manifolds and in the chamber are presented. The thrust–time profile and detonation velocity calculated from measured pressure peaks are shown. To confirm the performance of a rocket engine with rotating detonation as a high energy gas generator, a model of a simple engine was designed, built, and tested. In the tests, the model of the engine was connected to the dump tank. This solution enables different environmental conditions from a range of flight from 16 km altitude to sea level to be simulated. The obtained specific impulse for pressure in the chamber of max. 1.2 bar and a small nozzle expansion ratio of about 3.5 was close to 1,500 m/s.

Thermochemical erosion analysis for graphite/carbon-carbon rocket nozzles

Abstract: A study is conducted to predict graphite/carbon–carbon nozzle erosion behavior in solid rocket motors for wide variations of propellant formulations. The numerical model considers the solution of Reynolds-averaged Navier– Stokes equations in the nozzle, heterogeneous chemical reactions at the nozzle surface, variable transport and thermodynamic properties, andheat conduction in the nozzlematerial. Twodifferent ablationmodels are considered and compared: a surface equilibriumapproach and afinite-ratemodel. Results show that the erosion rate is diffusion limited for metallized propellants, ensuring sufficiently high wall temperatures, and it is kinetic limited for nonmetallized propellants. For low surface temperatures, the twomodels are consistent with each other and predict the same erosion rate, while the surface equilibrium model overpredicts the recession at low surface temperatures. The calculated results show an excellent agreement with the experimental data from the ballistic test and evaluation system motor firings, and the finite-rate model actually improves the predictions when the kinetic-limited regime is approached.

Handbook of X-ray Astronomy

Abstract: X-ray astronomy was born in the aftermath of World War II as military rockets were repurposed to lift radiation detectors above the atmosphere for a few minutes at a time. These early flights detected and studied X-ray emission from the Solar corona. The first sources beyond the Solar System were detected during a rocket flight in 1962 by a team headed by Riccardo Giaccom at American Science and Engineering, a company founded by physicists from MIT. The rocket used Geiger counters with a system designed to reduce non-X-ray backgrounds and collimators limiting the region of sky seen by the counters. As the rocket spun, the field of view (FOV) happened to pass over what was later found to be the brightest non-Solar X-ray source; later designated See X-1. It also detected a uniform background glow which could not be resolved into individual sources. A follow-up campaign using X-ray detectors with better spatial resolution and optical telescopes identified See X-1 as an interacting binary with a compact (neutron star) primary. This success led to further suborbital rocket flights by a number of groups. More X-ray binaries were discovered, as well as X-ray emission from supernova remnants, the radio galaxies M87 and Cygnus-A, and the Coma cluster. Detectors were improved and Geiger counters were replaced by proportional counters, which provided information about energy spectra of the sources. A constant challenge was determining precise positions of sources as only collimators were available.

VX-200 Magnetoplasma Thruster Performance Results Exceeding Fifty-Percent Thruster Efficiency

Abstract: H IGH-POWER electric propulsion thrusters can reduce propellant mass for heavy-payload orbit-raising missions and cargo missions to the moon and near-Earth asteroids, and they can reduce the trip time of robotic and piloted planetary missions [1–4]. The Variable Specific Impulse Magnetoplasma Rocket (VASIMR®) VX-200 engine is an electric propulsion system capable of processing power densities on the order of 6 MW=m with a high specific impulse (4000 to 6000 s) and an inherent capability to vary the thrust and specific impulse at a constant power. The potential for a long lifetime is due primarily to the radial magnetic confinement of both ions and electrons in a quasi-neutral flowing plasma stream, which acts to significantly reduce the plasma impingement on the walls of the rocket core. High-temperature ceramic plasma-facing surfaces handle the thermal radiation: the principal heat transfer mechanism from the discharge. The rocket uses a helicon plasma source [5,6] for efficient plasma production in the first stage. This plasma is energized further by an ion cyclotron heating (ICH) RF stage that uses left-hand polarized slow-mode waves launched from the high field side of the ion cyclotron resonance. Useful thrust is produced as the plasma accelerates in an expanding magnetic field: a process described by conservation of the first adiabatic invariant as the magnetic field strength decreases in the exhaust region of the VASIMR [7–9]. End-to-end testing of the VX-200 engine has been undertaken with an optimum magnetic field and in a vacuum facility with sufficient volume and pumping to permit exhaust plumemeasurements at low background pressures. Experimental results are presented with the VX-200 engine installed in a 150 m vacuum chamber with an operating pressure below 1 10 2 Pa (1 10 4 torr), and with an exhaust plume diagnostic measurement range of 5 m in the axial direction and 1 m in the radial directions. Measurements of plasma flux, RF power, and neutral argon gas flow rate, combined with knowledge of the kinetic energy of the ions leaving the VX-200 engine, are used to determine the ionization cost of the argon plasma. A plasmamomentum flux sensor (PMFS)measures the force density as a function of radial and axial positions in the exhaust plume. New experimental data on ionization cost, exhaust plume expansion angle, thruster efficiency, and total force are presented that characterize the VX-200 engine performance above 100 kW. A semiempirical model of the thruster efficiency as a function of specific impulse has been developed to fit the experimental data, and an extrapolation to 200 kWdc input power yields a thruster efficiency of 61% at a specific impulse of 4800 s.

Analysis of Curved-Cooling-Channel Flow and Heat Transfer in Rocket Engines

Abstract: DOI: 10.2514/1.B34163 Coolant-flow modeling in regeneratively cooled rocket engines fed with turbomachinery is a challenging task because of the high wall-temperature gradient, the high Reynolds number, the high aspect ratio of the channel cross section,andthecurvedgeometry.Inthepresentstudy,tobettercomprehendtheroleofthethrust-chambershapeof a rocket engine on the heat exchange, computations of supercritical hydrogen flow in single- and double-curvature channels are carried out. In particular, a parametric numerical analysis of the flow in an asymmetrically heated rectangular channel with a high aspect ratio and various radii of curvature is performed by means of a Reynoldsaveraged Navier–Stokes solver for real fluids, which is validated against experimental data of heated and curvedchannel flow taken from open literature. Results permit the effect of curvature on global heat transfer coefficient, pressure loss, and bulk temperature increase to be quantified.

Side Loads in Subscale Dual Bell Nozzles

Abstract: TODAY’S European heavy lifter Ariane 5 features a parallel staged design, where a cryogenic main stage is supported by two solid boosters generating the main part of the liftoff thrust Its original objectivewas to deliver heavy payloads to a low Earth orbit Nowadays Ariane 5’s dual GTO payload capability is in focus In opposition to tandem-staged rocket systems, like Ariane 4, the main stage engine Vulcain 2 has to be ignited on the ground for security reasons to assure proper running before solid boosters’ ignition and rocket takeoff Because of this design concept, the main stage engine has to fulfill a wide range of operation conditions, from sea level to near vacuum To reduce undesired side loads that would affect the engine, the rocket structure, and even the payload itself, the nozzle area ratio is limited, preventing flow separation at sea level This area ratio limitation leads to performance losses as the engine’s exhaust flow is driven overexpanded at sea level and highly under expanded at high altitudes To optimize the overall Isp of an engine during ascent, the use of altitude-adaptive nozzles, where the thrust generation is not only optimized at one specific altitude, comes into focus as the subsystem with the most promising performance gain Different concepts were developed to circumvent the limitation in area ratio of conventional nozzles The commonly discussed solutions are plug, extendible, and dual bell nozzles The characteristic contour inflection of the dual bell nozzle divides the nozzle into base and extension (Fig 1) and offers a one-step altitude adaptation At sea level, the contour inflection forces the flow to separate controlled and symmetrically (Fig 2) The base nozzle flows full and the extension is separated: the dual bell is operating in sea levelmode Because of a smaller effective area ratio the sea level Isp increases compared with a conventional nozzle (Fig 3) At the designed altitude theflow attaches abruptly to thewall of the extension down to the exit plane (Fig 4) This transition to high-altitude mode results in a short time Isp loss but later on in a higher vacuum performance The dual bell’s major advantage is the absence of anymoving parts Only minor changes to the design and the structure of already operating rocket engines would be necessary The concept of applying a contour inflection was first mentioned by Foster and Cowles [1] within a study on flow separation in supersonic nozzles Various solutions were suggested to prevent uncontrolled flow separation The onewith an inflection dividing the nozzle in two parts was later patented as the dual bell nozzle by Rocketdyne in 1968 The first experimental study was performed by Horn and Fisher [2] with different extension contour design approaches in cold flow subscale tests The transition from one operating mode to the other is particularly of interest as the flow potentially separates asymmetrically within the extension, resulting in a strong side load peak The dual bell topic was introduced in the late 90s into Europe’s community [3] Hagemann et al [4] presented in 2000 experimental cold as well as hot flow studies with respect to side load generation One remarkable fact is that the side load peak during retransition (while the nozzle is shut down) was shown to be significantly higher than during transition An opposite result is given in studies performed since (eg, by Hieu et al [5]) where the transition to highaltitude mode generates higher side loads The experimental cold flow results [4] were recalculated at DLR, German Aerospace Center by Karl and Hannemann [6] using the inhouse code TAU The transient simulations showed that the calculated side load peak during transition mainly depended on the nozzle Presented as Paper 2010-6729 at the 46th AIAA Joint Propulsion Conference, Nashville, TN, 25–282010; received 3November 2010; revision received 3 February 2011; accepted for publication 8 February 2011 Copyright © 2011 by DLR, German Aerospace Center Published by the American Institute of Aeronautics and Astronautics, Inc, with permission Copies of this paper may be made for personal or internal use, on condition that the copier pay the $1000 per-copy fee to theCopyright Clearance Center, Inc, 222RosewoodDrive, Danvers,MA01923; include the code 0748-4658/ 11 and $1000 in correspondence with the CCC Research Scientist, Institute of Space Propulsion, Langer Ground Head of Nozzle Group, Institute of Space Propulsion, Langer Ground JOURNAL OF PROPULSION AND POWER Vol 27, No 4, July–August 2011

Unmanned aerial vehicle system

Abstract: A UAV includes: a rocket body, having a rocket motor and a payload section; a parachute coupled with the payload section; an image capture device; a magnetometer to provide a compass reference for images taken from the image capture device; and a transmitter to communicate image and compass data to a remote receiver. Compass bearings are overlaid on image data from the image capture device. A handheld launch unit includes an ignition system, having an activation mechanism and an igniter to activate the rocket motor. A safety pin prevents electrical current from flowing to the igniter until the pin is removed. An accelerometer and/or magnetometer determines an angular orientation of the UAV. Software verifies that the angle is within a user-defined safety limit before activating the igniter.

Experimental Study on Spray Characteristics of Gas-Centered Swirl Coaxial Injectors

Abstract: Gas-centered swirl coaxial injectors have become an important subject of study for staged combustion rocket engines with hydrocarbon fuels. While these injectors are employed successfully in rocket engines, it is very rare to find the related research results as applicable to design data. An experimental study on spray characteristics of gas-centered swirl coaxial injectors has been performed. The effects of momentum flux ratio and recess length on the spray characteristics have been investigated by cold flow tests with a photographic technique. The liquid intact length L, which profoundly affects the global spray characteristics, decreases as the momentum flux ratio M increases. The critical momentum flux ratio Mc is introduced to identify the flow patterns as internal or external mixing in the injectors. Concerning the effect of the recess length lR, it is shown that the spray cone angle and the drop size decrease as lR increases.

Test-Based Thermal Decomposition Simulation of AP/HTPB and AP/HTPE Propellants

Abstract: The ammonium perchlorate composite propellant is a common choice for solid rocket propulsion. The externally heated rocket via fires, for instance, can cause the energetic substance to ignite, and this may lead to a thermal runaway event or a slow cook-off phenomenon marked by a severe explosion. To develop preventive measures to reduce the possibility of such accidents in propulsion systems, we investigate the ignition and initiation characteristics of ammonium-perchlorate-based propellants andperform the thermal decomposition simulation of a thermally loaded solid rocket motor.

Numerical Simulation of Liquid Rocket Engine Thrust Chamber Regenerative Cooling

Abstract: The objective of this paper is to develop and compare multidisciplinary, multidimensional, numerical methodologies to predict the hot-gas-side and coolant-side heat transfer in regeneratively-cooled rocket engine thrust chamber. The first methodology used empirical equations to simulate the hot-gas convective and radiative heat transfer; the second methodology used computational fluid dynamics to simulate the hot-gas convective and radiative heat transfer; heat transfer in the coolant and in the cooling channel was solved in a conjugate manner for the two methodologies. Systematic parametric studies on effects of combustion chemistry, radiation coupling, and grid refinement were performed and assessed. The methodologies were assessed by existing data from Arnold Engineering Development Center high-enthalpy nozzle tests and hot-firing test of a LO 2 -LH 2 thrust chamber. Results indicate that the second methodology with finite-rate chemistry employed in this study can be an effective method for predicting the flow and heat transfer in regeneratively-cooled thrust chamber, but needs further modifications.

High-repetition-rate PIV investigations on a generic rocket model in sub- and supersonic flows

Abstract: High-repetition-rate PIV measurements were performed in the trisonic wind tunnel facility at the Bundeswehr University Munich in order to investigate the boundary layer parameters on a generic rocket model and the recirculation area in the wake of the model at Mach numbers up to Mach = 26 The data are required for the validation of unsteady flow simulations Because of the limited run time of the blow-down wind tunnel, a high-repetition-rate PIV system was applied to obtain the flow statistics with high accuracy The results demonstrate this method’s potential to resolve small-scale flow phenomena over a wide field of view in a large Mach number range but also show its limitations for the investigations of wall-bounded flows

Time-resolved burning of solid fuels for hybrid rocket propulsion

Abstract: Features such as lowcost, safety, throttleability, and a wide range of appealing applications (e. g., interplanetary landers, boosters for space launcher, upper stage for Vega launcher) make hybrid rocket engines a very attractive option for aerospace propulsion. However, problems such as low regression rates of the solid fuel and low combustion e©ciency have so far hindered the development of large-scale hybrid rocket engines. Space Propulsion Laboratory (SPLab) at Politecnico di Milano has developed a series of proprietary techniques to evaluate, on a relative grading, the quasi-steady regression rates of solid fuels while visualizing at the same time the §ame structure. Numerical modeling, thermochemical calculations, and mechanical testing complete the range of tools set up to assess the quality of new solid fuels. In this paper, HTPB polymer has been taken as baseline and characterized at laboratory level.

Characterization of the Performance of Paraffin / LiAlH4 Solid Fuels in a Hybrid Rocket System

Abstract: This investigation examined the burning characteristics of paraffin-based solid-fuel grains doped with various additive percentages (up to 28%) of lithium aluminum hydride (LiAlH4). In addition, the test sequence included examination of a paraffin-wax based fuel formulation containing 10% triethylaluminum and another formulation containing 10% diisobutylaluminum hydride. The fuel grains were cast into paper phenolic tubes and then tested in a cartridge-loaded hybrid rocket system. It was found that under similar test conditions, increased LiAlH4 additive increased the overall chamber pressure throughout the duration of the test, caused by an increase in the ratio of flame temperature to the molecular weight of the products. Due to deposits of unburned and unreacted fuel in downstream sections of the hybrid rocket motor, an accurate correlation between increased additive percentage and regression rate was not able to be found. It was determined that a new set of fuel grain formulations with changes to the overall fuel matrix (e.g., higher melting point wax) and/or changes to the energetic additive particles (e.g., reduced particle size) will allow for more accurate regression rate calculations and more favorable combustion characteristics. Despite the necessary modifications to the fuel formulations, the results from this series of tests showed that nearly all these solid-fuel formulations burned similarly. Qualitative comparisons of each type of fuel formulation proved to be a beneficial method for improving the solid-fuel formulations for future tests for hybrid rocket motor applications.

Color Schlieren imaging of high-pressure overexpanded planar nozzle flow using a simple, low-cost test apparatus

Abstract: Complex flow features within rocket nozzles can exert significant influence on both the dynamics and safety of rockets during flight. Specifically, under over-expanded flow conditions, during, low altitude flight, random, often large side loads can appear within nozzles. While significant research has focused on this classical problem, due to the high nozzle pressure ratios (NPR) extant across rocket nozzles, most experimental work: (1) has focused on measuring wall pressure distributions under conditions when side loads appear, (2) has been carried out in large government or industrial test facilities, and (3) has only provided limited, though crucially important, visualization data. This short paper describes the construction and operation of a very simple, low cost test apparatus that allows imaging of flow features within planar nozzles, under the high NPR conditions characteristic of medium-to-large rockets. Representative color Schlieren images of flow shock structure obtained within the test apparatus are also presented and briefly described.

Performance Modeling of a Paraffin Wax / Nitrous Oxide Hybrid Rocket Motor

Abstract: This paper describes the development of a one-dimensional, non-steady flow predictive performance model for hybrid rocket motors. The analysis relies on basic thermodynamic and gas dynamic assumptions. The analytical model was programmed in MATLAB and uses the NASA-CEA equilibrium chemistry code to determine gas combustion properties. An empirical regression rate correlation is employed in the model with ballistic coefficients found in the literature. The predictive model implements conservation of mass and the ideal gas equation of state to compute changes in chamber pressure. Instantaneous performance parameters such as theoretical thrust and specific impulse can be analyzed for different operating conditions defined by the user. To validate the analytical model, a laboratory-scale motor has been designed and manufactured using a paraffin wax and nitrous oxide propellant combination. Hot-fire test data are compared with the predictive modeling.

Optical Emissions Observed During the Charged Aerosol Release Experiment (CARE I) in the Ionosphere

Abstract: The in-flight engine firing of solid rocket motors in the ionosphere produces an artificial dusty plasma. Optical emissions of sunlight scattered from the dust particles yield measurements of the dust location and flow velocities. Charging by ambient ionospheric electrons of the particulates yields dust particles that stream across the magnetic field lines. These exhaust particles initiate plasma turbulence in the ionosphere that can scatter radar waves. If the exhaust cloud itself passes over in situ particle or plasma wave detectors, measurements can be made of increased dusty plasma wave turbulence and plasma densities. To demonstrate long-range detection of rocket engine burns in the ionosphere, the Charged Aerosol Release Experiment (CARE I) was conducted in September 2009. Optical observations from CARE I provided measurements of both the dust particle distributions and the interactions of the molecular component of the rocket exhaust in the ionosphere.

CFD Simulation of a Hybrid Rocket Motor with Liquid Injection

Abstract: In this paper, a commercial CFD code is used to simulate different hybrid rocket motor configurations applying liquid N2O as the oxidizer and paraffin wax as the fuel. This work is the prosecution of a previous study performed to simulate hybrid rockets with diaphragms of different geometries placed inside the combustion chamber, where N2O was injected in gaseous phase, instead of using liquid. With respect to the previous study, liquid injection has been introduced, together with the droplets vaporization inside the combustion chamber and their full coupling with the eulerian gas phase, in terms of both heat exchange and momentum exchange. The main objective is the description of the proper numerical models to be applied in test cases in which liquid injection has to be represented. The most important differences with respect to the simulations where only gas is injected are also discussed. In order to validate the CFD output, experimental results coming from two different design scales are used: a laboratory scale and an increased scale. For each of these two scales, different rocket configurations and geometries have been studied. The different geometries studied include: a lab-scale rocket with a cylindrical grain and with a 4-hole diaphragm inserted at the 24% of the grain length, a 1-hole diaphragm lab-scale motor and an increased-scale hybrid rocket with a 1-hole diaphragm and without any diaphragm. For each test case, a comparison with the related experiment is presented and discussed. The simulations have been run in steady state conditions, with simplified chemical reactions, liquid oxidizer injection and no paraffin entrainment. The simulations show a good agreement with the experimental results of the different rocket configurations analyzed: the maximum error on efficiency is 7%. The CFD predicts (both in the case of gas and liquid injection) a higher efficiency for the rocket geometries provided with a diaphragm with respect to the same geometries without a mixing device and this is in accord with experiments. CFD results also show some peculiar phenomena about liquid injection.

Rocket Based Combined Cycle Hypersonic Vehicle Design for Orbital Access

Abstract: Several rocket boosted RBCC (Rocket-Based Combined Cycle) designs were created and analyzed for suitability for an access-to-space TSTO vehicle system. Level 1 analysis was performed that included vehicle closure, weight breakdown (to component level), flight trajectory data, propulsion and aerodynamic performance. Various inlet shapes were considered, such as symmetric and non-symmetric inlets. Also, several geometric configurations were studied, such as single flowpath, dual flowpath, engine-on-top and engine-on-bottom. At the end of the study, the number of candidate designs were reduced to two; one as a primary design and the other as the backup design. The primary design, a dual flowpath, engine-on-top design was selected for further analysis. Due to the large volume of the payload, the dual flowpath design was found to be more suitable than a single flowpath design. The vehicle reentry performance was analyzed. Basic cost estimation summary is included exploring the potential for commercial use.

Use of a Microphone Phased Array to Determine Noise Sources in a Rocket Plume

Abstract: A 70-element microphone phased array was used to identify noise sources in the plume of a solid rocket motor. An environment chamber was built and other precautions were taken to protect the sensitive condenser microphones from rain, thunderstorms and other environmental elements during prolonged stay in the outdoor test stand. A camera mounted at the center of the array was used to photograph the plume. In the first phase of the study the array was placed in an anechoic chamber for calibration, and validation of the indigenous Matlab(R) based beamform software. It was found that the "advanced" beamform methods, such as CLEAN-SC was partially successful in identifying speaker sources placed closer than the Rayleigh criteria. To participate in the field test all equipments were shipped to NASA Marshal Space Flight Center, where the elements of the array hardware were rebuilt around the test stand. The sensitive amplifiers and the data acquisition hardware were placed in a safe basement, and 100m long cables were used to connect the microphones, Kulites and the camera. The array chamber and the microphones were found to withstand the environmental elements as well as the shaking from the rocket plume generated noise. The beamform map was superimposed on a photo of the rocket plume to readily identify the source distribution. It was found that the plume made an exceptionally long, >30 diameter, noise source over a large frequency range. The shock pattern created spatial modulation of the noise source. Interestingly, the concrete pad of the horizontal test stand was found to be a good acoustic reflector: the beamform map showed two distinct source distributions- the plume and its reflection on the pad. The array was found to be most effective in the frequency range of 2kHz to 10kHz. As expected, the classical beamform method excessively smeared the noise sources at lower frequencies and produced excessive side-lobes at higher frequencies. The "advanced" beamform routine CLEAN-SC created a series of lumped sources which may be unphysical. We believe that the present effort is the first-ever attempt to directly measure noise source distribution in a rocket plume.

Measurements of Density and Speed of Sound of JP-10 and a Comparison to Rocket Propellants and Jet Fuels

Abstract: Densities of the missile fuel JP-10 were measured with two vibrating-tube densimeters. The combined range of the data is from 270 to 470 K, with pressures up to 30 MPa. The speed of sound in the fuel was measured with a propagation time method at ambient pressure from 278.15 to 343.15 K. The results for density and speed of sound at ambient pressure were combined to obtain the adiabatic compressibility. Correlations are reported that represent the temperature and pressure dependence of the experimental density data within their estimated uncertainty. The properties of JP-10 are compared to those of other previously measured jet and rocket fuels.

Numerical and Experimental Investigation on Vortex Injection in Hybrid Rocket Motors

Abstract: A study on vortex injection in hybrid rocket motors with nitrous oxide as the oxidizer and paraffin as the fuel has been performed. The investigation followed two paths: first of all, the flow field was simulated with a CFD code, and then burn tests were performed on a lab-scale rocket. The CFD analysis had the dual purpose to help the design of the lab motor and to understand the physics underlying the vortex flow coupled with the combustion process. Numerical analysis was focused on the comparison with axial injection. Vortex injection produces a more diffuse flame in the combustion chamber and improves the mixing process of the reactants, both aspects concurring to increase the efficiency of the motor. A helical streamline develops downstream the injection region, and the pitch is highly influenced by combustion, that tends to straighten the flow due to the acceleration imposed by the temperature rise to the axial velocity component. The tangential velocity, on the contrary, is far less influenced by this effect. Experimental tests with the same chamber geometry have been performed with both pressurized and self-pressurized oxidizer. Measured performances showed an increase in regression rate up to 51% and a combustion efficiency that rises from values lower than 80% in the axial injection configuration up to more than 90% with vortex. Moreover, a reduction of the instabilities in the chamber pressure has been measured. Issues requiring further investigation concern the motor exhausts: both experimental and numerical analyses showed that there is a residual tangential velocity component in the plume; this, coupled with a noise suppressor system downstream the nozzle in the test apparatus, showed severe instabilities in the vortex configuration thrust measurements, not reported in chamber pressure burn data and not affecting axial injection.

Intermetallic Compounds as Fuels for Composite Rocket Propellants

Abstract: Aluminized composite propellants have long suffered from efficiency and thermal challenges related to production of condensed phase slag droplets during operation. In an effort to mitigate the production of large droplets, a mechanically activated intermetallic forming nickel-aluminum compound was substituted for a portion of a propellant’s aluminum fuel. The resulting agglomerate size and burning rate of this propellant was compared to a standard aluminized AP/HTPB propellant. Addition of mechanically activated fuel particles increased the burning rate exponent of the propellant, while simultaneously decreasing condensed phase agglomerate size from 235 µm (for the control propellant) to 90 µ m( for the propellant containing 75 wt.% Ni-Al fuel). As such, intermetallic forming fuels may provide a route for increasing efficiency in solid rocket motors by simultaneously reducing the need for burning rate catalysts and minimizing two-phase nozzle flow losses.

Space-Time Accuracy Assessment of CFD Simulations for the Launch Environment

Abstract: Time-accurate high-fidelity Computational Fluid Dynamics (CFD) simulations of the launch environment are an important part of the successful launch of new and existing space vehicles. The capability to accurately predict certain aspects of the launch environment, such as ignition overpressure (IOP) waves and launch acoustics, is paramount to mission success. Implicit dual-time stepping methods represent one approach to provide accurate computational results in a timely manner. Two simplified test cases related to the launch environment are examined. The first test case models the IOP waves generated from a 2D planar jet located above a 45-degree flat plate, while the second case investigates launch acoustic noise generated from the jet of a rocket impinging on an axisymmetric flame trench and mobile launcher. Sensitivity analysis has been performed and a verification procedure was applied to investigate the necessary spatial and temporal resolution requirements for CFD simulations of the launch environment using an implicit dual-time method.

A Study on Combustion Efficiency of a Paraffin-based Hybrid Rockets

Abstract: A hybrid rocket has been considered to be a new hopeful rocket engine for safety, low cost, less pollution and so on. However, there are some technical problems. Those are low regression rate, low combustion efficiency and so on. Especially, low regression rete is key problem. We have been studying paraffin-based fuel which burn 3 to 5 times faster than that of HTPB. The reason why paraffin-based fuel has high regression rate is that this regression proceeds almost without vapourization. Therefore, liquid fuel is supplied to frame zone. This phenomenon makes combustion efficiency of paraffin-based hybrid rockets low level. We consider that it is effective to set up baffle plate inside combustion chamber and post combustion chamber behind baffle plate for atomization of the liquid fuel.