Showing papers on "Rocket published in 2015"

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Abstract: At this point in its development, the solar sail can be characterized as fairly late in its theoretical phase and fairly early in its developmental phase. It is probably equivalent to the chemical rocket in 1930, the automobile in 1900, and the heavier-than-air aircraft in 1910.

Approach for Modeling Rocket Plume Impingement and Dust Dispersal on the Moon

Abstract: When a lander approaches the lunar surface, the plume from the descent engine impinges on the ground and entrains loose regolith into a high-velocity spray. This problem is simulated with a hybrid ...

Simulation of Gaseous Oxygen/Hydroxyl-Terminated Polybutadiene Hybrid Rocket Flowfields and Comparison with Experiments

Abstract: Numerical simulations of the flowfield in a gaseous oxygen/hydroxyl-terminated polybutadiene hybrid rocket engine are carried out with a Reynolds-averaged Navier–Stokes solver including detailed gas/surface interaction modeling based on surface mass and energy balances. Fuel pyrolysis is modeled via finite-rate Arrhenius kinetics. A simplified two-step global reaction mechanism is considered for the gas-phase chemistry to model the combustion of 1,3-butadiene in oxygen. Results are compared with the firing test data obtained from a laboratory-scale hybrid rocket in which gaseous oxygen is fed into axisymmetric hydroxyl-terminated polybutadiene cylindrical grains through an axial conical subsonic nozzle. With the oxidizer fed by this injector, which generates nonuniform conditions at the entrance of the fuel port, the fuel regression rate is shown to increase several times with respect to the case of homogeneous injection of the oxidizer through all the grain port area, in agreement with the experimental f...

Comparative Evaluation Between Experiment and Simulation for a Transverse Instability

Abstract: A transverse combustion instability study composed of experimental and hybrid Reynolds-averaged Navier–Stokes/large eddy simulation results is described. The study is focused on gas-centered swirl coaxial injector elements like those used in the main chambers of oxidizer-rich staged-combustion engines. Experimental results are obtained from a self-excited, multi-element, high-pressure model rocket combustor. The simulation uses a velocity forcing technique to match the frequency and amplitude of the pressure oscillations that are measured in the combustor. Chemiluminescence is used to indirectly measure heat release rate in the experimental investigation. Detailed comparisons between the predicted and measured pressure field and the predicted and measured heat release modes are made. Modal decomposition is used to systematically compare the heat release modes. Results show successful computational replication of the experimental unsteady environment through the use of the velocity forcing technique and th...

Longitudinal Control Strategy for Hypersonic Accelerating Vehicles

Abstract: Recent work on small scale access to space systems found that by incorporating a scramjet powered stage into the launch stack the percent payload into orbit can be improved significantly over conventional rocket only systems. However, in order to gain the most out of the airbreathing stage the flight dynamic pressure across the accelerating trajectory becomes an important design consideration. If it is too high then the structural loading and aerodynamic heating can be excessive, if it is too low scramjet performance will suffer and unreasonably large wings will be required to generate enough lift. A common compromise is to fly a constant dynamic pressure trajectory. This paper presents and compares three dynamic pressure controller methodologies that were developed for a scaled down version of the winged cone vehicle developed by NASA during the 90s. The controllers investigated are a constant gain PID, a gain scheduling PID and a pole placement PI. The performances of the controllers are demonstrated along a 50kPa dynamic pressure accelerating trajectory. It is shown that a unique criteria that must be satisfied is an initial flight path angle that satisfies a zero rate of change in dynamic pressure, and an underlying angle of attack for ‘lift equals weight’.

Boron and Polytetrafluoroethylene as a Fuel Composition for Hybrid Rocket Applications

Abstract: A composition consisting of 80% polytetrafluoroethylene and 20% boron (by weight) was considered as a potential high-density solid fuel mixture for mixed hybrid rocket propulsive applications. Constant-pressure strand burner experiments for the given formulation indicated a low-pressure self-deflagration limit of approximately 2.2 MPa (319 psia), and a burning rate correlation rb[cm/s]=0.042(P[MPa])0.531 was determined. Pressurized counterflow burner experiments conducted using pure oxygen revealed formation of surface char, which prevented measurement of solid fuel regression rates below 2 MPa, indicating an additional resistance for heat and mass transfer. Static-fired rocket motor experiments, conducted to determine the pressure and flow dependencies of the system, exhibited characteristic exhaust velocity efficiencies ranging from approximately 86 to 96%. Whereas classical hybrids do not have a strong dependence of fuel regression rate on pressure, a pressure dependence was observed in this system bel...

Influence of Cooling-Gas Properties on Film-Cooling Effectiveness in Supersonic Flow

Abstract: Film cooling is considered a prerequisite for the safe operation of future high-performance rocket engines Wall-normal or inclined cooling-gas injection into a laminar boundary-layer airflow at Ma

Propellant Injector Influence on Liquid-Propellant Rocket Engine Instability

Abstract: The avoidance of acoustic instabilities, which may cause catastrophic failure, is demanded for liquid-propellant rocket engines. This occurs when the energy released by combustion amplifies acoustic disturbances; it is therefore essential to avoid such positive feedback. Although the energy addition mechanism operates in the combustion chamber, the propellant injector system may also have considerable influence on the stability characteristics of the overall system, with pressure disturbances in the combustion chamber propagating back and forth in the propellant injectorchannels.Theintroducedtimedelaymayaffectstability,dependingontheratioofthewavepropagationtime throughtheinjectortotheperiodofthecombustionchambersacousticmodes.Thisstudyfocusesontransverse-wave liquid-propellant rocket engine instabilities using a two-dimensional polar coordinate solver (with averaging in the axial direction) coupled to one-dimensional solutions in each of the coaxial oxygen–methane injectors. A blockage in one (or more) of the injectors is analyzed as a stochastic event that may cause an instability. A properly designed temporaryblockage of oneor more injectors can also be used for control of an oscillation introduced by any physical event.Thestochasticanddesignvariablesparameterspaceis exploredwiththepolynomial chaosexpansionmethod.

Theoretical and experimental analysis of liquid layer instability in hybrid rocket engines

Abstract: The combustion behavior of different hybrid rocket fuels has been analyzed in the frame of this research. Tests have been performed in a 2D slab burner configuration with windows on two sides. Four different liquefying paraffin-based fuels, hydroxyl terminated polybutadiene (HTPB) and high-density polyethylene (HDPE) have been tested in combination with gaseous oxygen (GOX). Experimental high-speed video data have been analyzed manually and with the proper orthogonal decomposition (POD) technique. Application of POD enables the recognition of the main structures of the flow field and the combustion flame appearing in the video data. These results include spatial and temporal analysis of the structures. For liquefying fuels these spatial values relate to the wavelengths associated to the Kelvin Helmholtz Instability (KHI). A theoretical long-wave solution of the KHI problem shows good agreement with the experimental results. Distinct frequencies found in the POD analysis can be related to the precombustion chamber configuration which can lead to vortex shedding phenomena.

Measurement of atomic oxygen in the middle atmosphere using solid electrolyte sensors and catalytic probes

Abstract: . The middle- and upper-atmospheric energy budget is largely dominated by reactions involving atomic oxygen (O). Modeling of these processes requires detailed knowledge about the distribution of this oxygen species. Understanding the mutual contributions of atomic oxygen and wave motions to the atmospheric heating is the main goal of the rocket project WADIS (WAve propagation and DISsipation in the middle atmosphere). It includes, amongst others, our instruments for the measurement of atomic oxygen that have both been developed with the aim of resolving density variations on small vertical scales along the trajectory. In this paper the instrument based on catalytic effects (PHLUX: Pyrometric Heat Flux Experiment) is introduced briefly. The experiment employing solid electrolyte sensors (FIPEX: Flux φ(Phi) Probe Experiment) is presented in detail. These sensors were laboratory calibrated using a microwave plasma as a source of atomic oxygen in combination with mass spectrometer reference measurements. The spectrometer was in turn calibrated for O with a method based on methane. In order to get insight into the horizontal variability, the rocket payload had instrument decks at both ends. Each housed several sensor heads measuring during both the up- and downleg of the trajectory. The WADIS project comprises two rocket flights during different geophysical conditions. Results from WADIS-1 are presented, which was successfully launched in June 2013 from the Andoya Space Center, Norway. FIPEX data were sampled at 100 Hz and yield atomic oxygen density profiles with a vertical resolution better than 9 m. This allows density variations to be studied on very small spatial scales. Numerical simulations of the flow field around the rocket were done at several points of the trajectory to assess the influence of aerodynamic effects on the measurement results. Density profiles peak at 3 × 1010 cm−3 at altitudes of 93.6 and 96 km for the up- and downleg, respectively.

Computational Fluid Dynamics Simulation of Hybrid Rockets of Different Scales

Abstract: CFX software is used to simulate different hybrid rocket configurations, applying liquid N2O as the oxidizer and paraffin as the fuel. This work is the prosecution of a previous paper analyzing liquid injection in a lab-scale hybrid rocket. It is focused on the formulation of the most suitable simulation technique to represent another type of liquid injector, compared with the one described in the previous paper. It also aims at extending the computational fluid dynamics simulation approach to hybrid rockets of larger scales. To validate computational fluid dynamics output, experimental results coming from both a laboratory scale and an increased-scale engine have been used. The different geometries studied include an increased-scale engine with a cylindrical grain having no diaphragm, the same rocket with a one-hole diaphragm inside the fuel grain, and a lab-scale rocket with a one-hole diaphragm. Simulations are steady state, and combustion derives from a single-phase chemical reaction. Liquid injection...

Computational Fluid Dynamics Simulations of a GO2/GH2 Single Element Combustor

Abstract: solution is obtained on a mesh with 1.5 million nodes. The complexity of the model is gradually increased until the model is capable of predicting the wall heat flux. The analysis of numerical results shows a significant effect of boundary conditions on wall heat flux predictions. The comparison of the Reynolds-averaged Navier–Stokes simulationswiththeexperimentaldatademonstratesthecapabilityofReynolds-averagedNavier–Stokessimulations to predict wall heat fluxes in a rocket combustion chamber.

Flight Testing of Terrain-Relative Navigation and Large-Divert Guidance on a VTVL Rocket

Abstract: Since 2011, the Autonomous Descent and Ascent Powered-Flight Testbed (ADAPT) has been used to demonstrate advanced descent and landing technologies onboard the Masten Space Systems (MSS) Xombie vertical-takeoff, vertical-landing suborbital rocket. The current instantiation of ADAPT is a stand-alone payload comprising sensing and avionics for terrain-relative navigation and fuel-optimal onboard planning of large divert trajectories, thus providing complete pin-point landing capabilities needed for planetary landers. To this end, ADAPT combines two technologies developed at JPL, the Lander Vision System (LVS), and the Guidance for Fuel Optimal Large Diverts (G-FOLD) software. This paper describes the integration and testing of LVS and G-FOLD in the ADAPT payload, culminating in two successful free flight demonstrations on the Xombie vehicle conducted in December 2014.

Propulsive jet influence on generic launcher base flow

Abstract: Afterbody flow phenomena represent a significant source of uncertainties in the design of a launcher. Therefore, there is a demand for measuring such flows in wind tunnels. For propulsive jet simulation a new jet facility was integrated into a hypersonic/supersonic wind tunnel. The jet simulation resembles the generic model of a staged rocket launcher. The design and the qualification of the facility are reported. This includes measurements of pressure, temperature and Mach number distribution. Pressure and Schlieren measurements are conducted in the wake of the generic launcher. The unsteady pressure characteristics at the generic rocket base and fairing are analyzed for supersonic and hypersonic freestream. The influence of the under-expanded jet is reported and the jet temperatures are varied. On the base fluctuations at a Strouhal number around 0.25 dominates supersonic freestream flows. Additionally, a fluctuation level increase on the base is observed for Strouhal numbers above 0.75 in hypersonic flow regime, which is attributed to the interactions of wake flow and jet.

Steady-state coupled analysis of flowfields and thermochemical erosion of C/C nozzles in hybrid rocket motors

Abstract: A hybrid rocket can be used in various applications and is an attractive propulsion system. However, serious erosion of nozzles is common in motor firing operations, which could restrict the application of hybrid rocket motors. Usually, the serious erosion is attributed to the high concentration of oxidizing species in hybrid motors, while the details of flowfields in the motors are not paid special attention to. In this paper, first the thermochemical erosion of C/C nozzle is simulated coupled with the flowfields in a 98% H2O2/hydroxyl-terminated polybutadiene (HTPB) hybrid rocket motor. The simulation is made on a typical axisymmetric motor, including a pre-combustion chamber, an aft-combustion chamber and nozzle structures. Thermochemical reactions of H2O, CO2, OH, O and O2 with C are taken into account. Second, the change of flowfields due to fuel regression during motor firing operations is considered. Nozzle erosion in different flowfields is evaluated. Third, the results of nozzle erosion in the coupled simulation are compared with those under uniform and chemical equilibrium flow and motor firing test results. The results of simulation and firing tests indicate that the thermochemical erosion of nozzles in hybrid motors should be calculated coupled with flowfields in the motor. In uniform and chemical equilibrium flowfields, the erosion rate is overestimated. The diffusion flame in hybrid motors protects the nozzle surface from the injected oxidizer and high temperature products in flowfields, leading to a relatively fuel-rich environment above the nozzle. The influence of OH and the geometry of motor should also be considered in the evaluation of nozzle erosion in hybrid motors.

R&D on Hydrocarbon-fueled RBCC Engines for a TSTO Launch Vehicle

Abstract: Combination of a air-breathing engines and rocket engines (termed as Rocket Based Combined Cycle engine) gives an opportunity to reduce onboard oxygen consumption and to increase system weight margins for various application from cruiser to launch vehicles. For launch vehicle applications, combination of a ramjet/scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines was proposed (termed as Rocket-Ramjet Combined Cycle engine, T. Kanda, et. al., JPP, 19(2003), 859), and related R&D activities are undergoing at Japan Aerospace Exploration Agency (JAXA), Kakuda Space Center targeting hydrogen as the fuel due to its high Isp and cooling performances. Use of hydrogen fuel, on the other hand, can be costly due to the fuel cost and difficult-to-handle nature of liquefied hydrogen. Thus, hydrocarbon fuel such as ethanol was under consideration for the application to the future reusable launch vehicles termed as the ‘reference system,’ targeting manned operation to LEO. This presentation is to show the status of system analysis, flight demonstration plan, and related research efforts.

Scaling of Hybrid Rocket Motors with Swirling Oxidizer Injection - Part 2

Abstract: The present research focuses on the use of a swirling oxidizer injection, in a hybrid motor with a classical cylindrical solid fuel grain with a circular port, ie a side burning grain The conducted study is mainly a numerical analysis, that pays a particular attention on the scaling behaviour of this particular configuration of hybrid rocket motor, in relation to the variation of the geometric swirl number and of the oxidizer mass flux