Showing papers on "Rocket published in 2014"
TL;DR: In this article, the results of computer code developing, verification and validation, making it possible to simulate unsteady processes of ignition and combustion of hydrogen fuel in rocket engines are developed.
418 citations
TL;DR: In this paper, a verification and validation of the chemical kinetic models for hydrogen oxidation were carried out through investigations on the ignition delay time on pressure, temperature, and equivalence ratio for hydrogen-oxygen mixtures.
148 citations
TL;DR: The High-Resolution Coronal Imager (Hi-C) was launched on a NASA sounding rocket on 11 July 2012 to investigate the dynamics of solar coronal structures at small spatial scales as mentioned in this paper.
Abstract: The High-Resolution Coronal Imager (Hi-C) was flown on a NASA sounding rocket on 11 July 2012. The goal of the Hi-C mission was to obtain high-resolution (≈ 0.3 – 0.4′′), high-cadence (≈ 5 seconds) images of a solar active region to investigate the dynamics of solar coronal structures at small spatial scales. The instrument consists of a normal-incidence telescope with the optics coated with multilayers to reflect a narrow wavelength range around 19.3 nm (including the Fe xii 19.5-nm spectral line) and a 4096×4096 camera with a plate scale of 0.1′′ pixel−1. The target of the Hi-C rocket flight was Active Region 11520. Hi-C obtained 37 full-frame images and 86 partial-frame images during the rocket flight. Analysis of the Hi-C data indicates the corona is structured on scales smaller than currently resolved by existing satellite missions.
138 citations
TL;DR: In this article, the effects of metal-based high-energy matter (Aluminum) into the content of the propellant produced within the scope of development project were investigated, and it was found that the burning rates and burning heat of new fuels manufactured by adding aluminum to the contents of the standard double base fuel (DB-1) increased.
91 citations
TL;DR: In this article, high-speed shadowgraph images show that the mechanism of core breakup and atomization differs between off-resonance and first transverse excitation conditions, with a core length reduction of up to 70% for conditions approaching those of naturally occurring high frequency combustion instabilities.
Abstract: Testing has been conducted with the BKH rocket combustor at the European Research and Technology Test Facility P8 for cryogenic rocket engines at DLR Lampoldshausen. BKH has multiple shear coaxial injectors and an exhaust modulation system for forcing excitation of acoustic resonances in the combustion chamber. Optical access windows allow the application of parallel high-speed shadowgraph and flame emission imaging of the near-injector region. This paper reports measurements of the intact liquid oxygen core during forced excitation of the first transverse acoustic mode. High-speed shadowgraph images show that the mechanism of core breakup and atomization differs between off-resonance and first transverse excitation conditions. The core length is found to decrease with increasing amplitude of acoustic pressure, or equivalently with transverse acoustic velocity, with a core length reduction of up to 70% for conditions approaching those of naturally occurring high frequency combustion instabilities. This de...
53 citations
TL;DR: In this article, the performance of an air turbo-rocket expander and a dual-mode ramjet was examined for a prospective supersonic cruise aircraft. But the performance was not optimal when the demand from the air-turbo-rocket matched the intake air flow capture and the engine was throttled varying the mixture ratio and the fan compression ratio.
39 citations
TL;DR: In this paper, the authors used a dense GPS network in China to track the ionospheric response to waves excited by the launch of the rocket that carried Shenzhou 10 spacecraft on 11 June 2013.
Abstract: We used a dense GPS network in China to track the ionospheric response to waves excited by the launch of the rocket that carried Shenzhou 10 spacecraft on 11 June 2013. The long-distance propagation of shock and acoustic waves were observed on both sides of the rocket's trajectory. On the southern side, the wave structures (characterized by a horizontal extension of ~1400 km and initial amplitudes of 0.3 total electron content unit (TECU) and 0.1 TECU for the shock and acoustic waves, respectively), traveled southwestward a distance of ~1500 km at mean velocities of 1011 m s−1 and 709 m s−1, respectively. On the northern side, northward propagating waves were seen to travel a distance of ~600 km with much smaller amplitudes of less than 0.05 TECU. Subsequent waves with amplitudes of less than 0.02 TECU could also be seen. Clear wave structures were found at a distance of ~600–2000 km from launch site.
38 citations
TL;DR: In this article, computational fluid dynamics applied to the analysis of a hybrid rocket motor with a diaphragm in the combustion chamber to enhance rocket performance has been described and compared.
Abstract: This paper describes computational fluid dynamics applied to the analysis of a hybrid rocket motor with a diaphragm in the combustion chamber to enhance rocket performance. This work follows the last author’s experiments: An engine was tested with nitrous oxide and paraffin wax as propellants. Several of the tests have been used as references for numerical simulations. The following approximations have been made: steady-state conditions, eddy dissipation model with one-step reaction, gaseous injection of fuel and oxidizer, and no droplets entrainment (typical of a paraffin grain). First of all, a single geometry without a diaphragm has been analyzed with different turbulence models (k-ω, k-ω shear stress transport, k-e, k-e renormalization group). It has been shown that the k-ω model predicts a lower flame temperature and chamber pressure than the k-e model. Then, five geometries have been studied to compare two different types of diaphragms (one hole and four holes) in two positions (24 and 33% of the to...
36 citations
TL;DR: In this article, a cooperative evolutionary method nested with an indirect approach is presented to perform the coupled optimization of hybrid rocket motor and trajectory for an upper stage, and a mission profile based on the Vega launcher is considered and the performance index is the payload inserted into the final orbit.
Abstract: Upper-stage motors used in small launchers constitute an application where hybrid rocket motors may be competitive. A coupled optimization of motor design and trajectory is needed for such an application due to mission characteristics and motor features. The present article presents a cooperative evolutionary method nested with an indirect approach to perform the coupled optimization of hybrid rocket motor and trajectory for an upper stage. The evolutionary method optimizes the parameters that affect the motor design (e.g., grain geometry) and feed system, whereas the indirect method optimizes the trajectory (i.e., thrust direction and motor switching times) for a given motor and mission. A mission profile based on the Vega launcher is considered and the performance index is the payload inserted into the final orbit. The hybrid rocket motor powers the third and last stage and has a pressurizing feed system that is partially regulated. The characteristics of the first and second solid rocket motor stages a...
35 citations
TL;DR: In this article, a Lyapunov-based nonlinear feedback control law is proposed to control the translational velocity vector and the attitude of the spacecraft, while attenuating the sloshing modes characterizing the internal dynamics.
34 citations
TL;DR: In this article, the instantaneous radiative forcing (RF) caused by major rocket engine emissions CO2, H2O, black carbon (BC), and Al2O3 (alumina) is estimated.
Abstract: Space transportation plays an important and growing role in Earth's economic system. Rockets uniquely emit gases and particles directly into the middle and upper atmosphere where exhaust from hundreds of launches accumulates, changing atmospheric radiation patterns. The instantaneous radiative forcing (RF) caused by major rocket engine emissions CO2, H2O, black carbon (BC), and Al2O3 (alumina) is estimated. Rocket CO2 and H2O emissions do not produce significant RF. BC and alumina emissions, under some scenarios, have the potential to produce significant RF. Absorption of solar flux by BC is likely the main RF source from rocket launches. In a new finding, alumina particles, previously thought to cool the Earth by scattering solar flux back to space, absorb outgoing terrestrial longwave radiation, resulting in net positive RF. With the caveat that BC and alumina microphysics are poorly constrained, we find that the present-day RF from rocket launches equals 16 ± 8 mW m−2. The relative contributions from BC, alumina, and H2O are 70%, 28%, and 2%. respectively. The pace of rocket launches is predicted to grow and space transport RF could become comparable to global aviation RF in coming decades. Improved understanding of rocket emission RF requires more sophisticated modeling and improved data describing particle microphysics.
TL;DR: In this paper, a liquid-fueled and valveless PDRE without the purge process was proposed, and a maximum operating frequency of 110 Hz was achieved when oxygen-enriched air with 25%∼45% oxygen by volume was employed.
TL;DR: In this article, the results of developing, verification and validation of mathematical model making it possible to simulate unsteady processes of ignition and combustion in rocket engines are presented and validated.
TL;DR: In this paper, the authors presented a modified optical scanning seeker with programmed and tracking controls with the simultaneous influence of external interferences from the direction of a self-guiding rocket.
TL;DR: In this paper, a tradeoff analysis is performed on a test case representative of the cooling system of a 1MN thrust class oxygen/hydrogen liquid rocket engine, where the authors find the channel aspect ratio that maximizes the heat extracted from the hot gas, for a given coolant pressure drop and hot-gas side wall temperature.
Abstract: A tradeoff analysis is performed on a test case representative of the cooling system of 1 MN thrust class oxygen/hydrogen liquid rocket engine. The aim of the analysis is to find the channel aspect ratio that maximizes the heat extracted from the hot gas, for a given coolant pressure drop and hot-gas side wall temperature. The analysis requires many cooling channel flow calculations that are performed by means of a simplified model, referred to as quasi-two-dimensional, and a three-dimensional conjugate heat transfer model based on numerical integration of the Navier–Stokes and Fourier’s equations. Both models are able to describe, with different level of details, the whole cooling device composed by the coolant and the solid domain, which is exposed to the hot gas. The fast quasi-two-dimensional approach is used to select channel geometries showing the same pressure loss. Discussion is made on results obtained with the more accurate three-dimensional model. Results identify, for the selected test case, a...
TL;DR: In this paper, the role of the longitudinal rib height on heat transfer enhancement is analyzed by means of a Reynolds-Averaged Navier-Stokes equations solver and a ribbed wall experiment is numerically reproduced to assess the capability of the simplified approach to properly capture the heat transfer enhancements.
Abstract: Heat transfer enhancement due to longitudinal wall ribs inserted in a rocket engine thrust chamber is analyzed by means of a Reynolds-Averaged Navier-Stokes equations solver. A ribbed wall experiment is numerically reproduced to assess the capability of the simplified approach to properly capture the heat transfer enhancement. Then, a parametric analysis on the role of the longitudinal rib height on heat transfer enhancement is made on a sample thrust chamber. Results show the expected heat increase related to the surface increase, and highlight the reduction of efficiency for increasing rib height due to the thermal stratification between ribs.
TL;DR: In this paper, the authors evaluate two rocket exhaust configurations within an Rocket-Based Combined Cycle engine model and show that the annular configuration yields an average Mach number 57% higher than the traditional circular configuration under conditions close to those for maximum air entrainment.
Abstract: Experimental tests are conducted to evaluate two rocket exhaust configurations within an Rocket-Based Combined Cycle engine model. Each configuration uses a single rocket chamber but expands the rocket exhaust to a different geometry. One configuration uses two traditional circular rocket nozzles whereas the other uses two annular exhaust geometries based on the Exchange Inlet design. Wall pressure data are collected along the length of each configuration to evaluate the air entrainment. The annular configuration is shown to increase the entrainment ratio by as much as 28% over the traditional circular rocket design. The annular configuration is also shown to choke the incoming airflow at a lower rocket chamber pressure more uniformly across the engine cross section. Total pressure measurements are also obtained at the engine exit plane, which show that the annular configuration yields an average Mach number 57% higher than the circular configuration under conditions close to those for maximum air entrain...
01 Jan 2014
TL;DR: In this paper, a multi-objective design optimization coupling evolutionary algorithms and trajectory optimization via pseudo-spectral methods has been conducted for the first stage of two-stage to orbit (TSTO) system with a rocket-based combined cycle (RBCC) engine which combines rockets and ramjets by blending two kinds of vehicle configurations with different aerodynamic characteristics.
Abstract: In this study, a multi-objective design optimization coupling evolutionary algorithms and trajectory optimization via pseudo-spectral methods has been conducted for the first stage of two-stage to orbit (TSTO) system with a rocket-based combined cycle (RBCC) engine which combines rockets and ramjets by blending two kinds of vehicle configurations with different aerodynamic characteristics. The design criteria include the minimization of fuel consumption and the maximization of the final Mach number up to a separation of the TSTO system at the maximum altitude under certain ranges of acceleration and dynamic pressure. The optimization results reveal a counteractive trend between the final Mach number and fuel mass ratio and the major impact of effective specific impulse on those two objectives, which is mainly controlled by thrust throttling parameter within the trajectory optimization. In addition, the RBCC-powered vehicle tends to fly at lower altitude to attain the minimum fuel mass ratio, in contrast to the case for maximum final Mach number, which is attributed to the hybrid aerodynamic performance of the two configurations. The insight gained here can be usefully applied to the design of high-performance RBCCpowered vehicles.
TL;DR: In this paper, two-way coupling and inertia relief methods are used to calculate the static deformations and aerodynamic characteristics of the deformed rocket. And the results highlight that the rocket deformation aspects are decided by the normal force distribution along the rocket length.
Abstract: The application and workflow of Computational Fluid Dynamics (CFD)/Computational Structure Dynamics (CSD) on solving the static aeroelastic problem of a slender rocket are introduced. To predict static aeroelastic behavior accurately, two-way coupling and inertia relief methods are used to calculate the static deformations and aerodynamic characteristics of the deformed rocket. The aerodynamic coefficients of rigid rocket are computed firstly and compared with the experimental data, which verified the accuracy of CFD output. The results of the analysis for elastic rocket in the nonspinning and spinning states are compared with the rigid ones. The results highlight that the rocket deformation aspects are decided by the normal force distribution along the rocket length. Rocket deformation becomes larger with increasing the flight angle of attack. Drag and lift force coefficients decrease and pitching moment coefficients increase due to rocket deformations, center of pressure location forwards, and stability of the rockets decreases. Accordingly, the flight trajectory may be affected by the change of these aerodynamic coefficients and stability.
TL;DR: In this article, the N2O and hydroxyl-terminated polybutadiene hybrid rocket motor was tested with a commercial off-the-shelf throttle valve and a solid rocket motor case adapted for hybrid rocket testing.
Abstract: Deep-throttle static test results from an N2O and hydroxyl-terminated polybutadiene hybrid rocket motor are presented. The nominal 800 N thrust level was turned down to less than 12 N while still maintaining stable and controlled combustion. This 67∶1 turndown was accomplished using a commercial off-the-shelf throttle valve and a solid rocket motor case adapted for hybrid rocket testing. During throttled motor tests, the pressure ratio across the injector grows from a nominal value of 2.0 to greater than 3.0. This feature contrasts with the observed behavior of liquid rockets, where the injector pressure ratio drops significantly during deep throttle. This characteristic likely supports the observed hybrid burn stability during deep throttle. Data comparisons with a physics-based, throttled, hybrid rocket burn model accurately match for combustor pressure, thrust, and propellant consumption. At throttle levels approaching 20% of nominal, the N2O exiting the throttle valve is entirely in a vapor state. The...
TL;DR: In this paper, an attempt was made to determine the regression rate of a hybrid fuel by using combustion chamber pressure, which was in good agreement with the results obtained using the weight loss method using the same motor and the same fuel and oxidizer combination.
28 Jul 2014
TL;DR: The Peregrine sounding rocket as discussed by the authors is a medium-scale liquid-fuel hybrid sounding rocket using storable propellants (paraffin wax and N2O) that will propel a 5 kg payload to the edge of space.
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 propel a 5 kg payload to the edge of space. This rocket, known as the Peregrine Sounding Rocket, 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. Results of ground testing performed using a heavy-weight configuration of the motor show that stable and efficient combustion has been achieved.
TL;DR: In this article, a closed-loop throttle controller for a laboratory-scale N2O and hydroxyl-terminated polybutadiene hybrid rocket motor is presented, which can significantly reduce the run-to-run burn variability typical of hybrid rocket motors.
Abstract: A closed-loop throttle controller for a laboratory-scale N2O and hydroxyl-terminated polybutadiene hybrid rocket motor is presented. Closed-loop throttling was achieved using commercial off-the-shelf valve hardware and a commercially available motor case adapted for hybrid rocket testing. Multiple and open- and closed-loop tests were performed to demonstrate that closed-loop control can significantly reduce the run-to-run burn variability typical of hybrid rocket motors. Closed-loop proportional/integral control algorithms featuring thrust or pressure feedback were used to track prescribed step and linear ramp profiles. Because the relationship between the selected throttle control valve position and the effective valve flow area was highly nonlinear, the effect of valve position on motor thrust/chamber pressure was measured open loop and curve fit to allow direct command of either total thrust or chamber pressure. Control law gains were tuned a priori using a numerical model and then adjusted using the a...
TL;DR: The Sharp Edge Flight Experiment II was launched from the Andoya rocket range in Norway on 22 June 2012, consisting of an extensively instrumented scientific payload on top of a two stage rocket configuration.
Abstract: The Sharp Edge Flight Experiment II was launched from the Andoya rocket range in Norway on 22 June 2012, consisting of an extensively instrumented scientific payload on top of a two stage rocket configuration. With an apogee of about 177 km, the vehicle achieved flight velocities up to 2790 m/s corresponding to Mach numbers up to 9.3. More than 96% of the pressure sensors provided excellent data during the ascent and descent phases. Measured pressure data from sensors at different locations of the scientific payload show consistent results concerning the aerodynamic behavior of the vehicle along the complete trajectory. Pressure fluctuations measured during flight show an excellent correlation to angle-of-attack variations. Calculated pressure coefficients from a computational fluid dynamics analysis at selected trajectory points are in good agreement with the measured pressure data.
TL;DR: In this paper, the authors present the overall strategy, the organization, and first experimental and numerical results of this joined effort to contribute to the development of improved hybrid propulsion systems, which aims to increase the scientific knowledge of the combustion processes in hybrid rockets using a strongly linked experimental-numerical approach.
13 Jan 2014
TL;DR: In this paper, a numerical code which uses the Large Eddy Simulation as a turbulent modeling and the Flamelet approach as combustion modeling is constructed, and the code is applied to the analysis of the swirling chamber.
Abstract: The objective of this study is to clarify inner state of a chamber of the Swirling-OxidizerFlow-Type Hybrid Rocket which is one of the types of a Hybrid Rocket by means of numerical fluid analysis. In this study, a numerical code which uses the Large Eddy Simulation as a turbulent modeling and the Flamelet approach as combustion modeling is constructed, and the code is applied to the analysis of the swirling chamber. On this occasion, in order to guarantee an applicability of the results, an experiment of the diffusion flame swirling burner is simulated by the code, and it is confirmed that results of the simulation are well corresponding qualitatively and partially quantitatively to experimental data. Then, a simulation for the chamber of the Swirling-Oxidizer-Flow-Type Hybrid Rocket is done by the numerical code, and it can be obtained that the numerical results are well corresponds qualitatively to visualized data of the experiment. Due to the analysis using this numerical code, structure of flow and flame, distributions of physical quantities and chemical species and state of turbulent eddies are clarified in the chamber of the hybrid rocket.
TL;DR: An algorithm is developed for determining the attitude of a spinning sounding rocket, able to track global positioning system (GPS) signals with intermittent availability but with enough accuracy to yield phase observables for the precise, three-axis attitude determination of a nutating rocket.
Abstract: An algorithm is developed for determining the attitude of a spinning sounding rocket. This algorithm is able to track global positioning system (GPS) signals with intermittent availability but with enough accuracy to yield phase observables for the precise, three-axis attitude determination of a nutating rocket. Raw, single-frequency GPS RF front-end data are processed by several filters to accomplish this task. First, a Levenberg-Marquardt algorithm (LMA) estimates GPS observables for multiple satellites by performing a least-squares fit to the accumulation outputs of a bank of correlators. These observables are then used as measurements in a Rauch-Tung-Striebel smoother that optimizes estimates of carrier phase, Doppler shift, and code phase. Finally, attitude determination is carried out by another batch filter that uses the single-differenced optimized carrier phase estimates between two antennas and an Euler dynamics model for the torque-free attitude motion of the spinning rocket. This second batch filter implements a combination of a substantially modified form of the LMA and the least-squares ambiguity decorrelation adjustment (LAMBDA) method. This design enables it to deal with integer ambiguities that change over long data gaps between times of carrier phase availability. The algorithm presented in this paper is applied to recorded RF data from a spinning sounding rocket mission to produce attitude quaternion and spin-rate estimates using a pair of antennas separated by a 0.3-m baseline. These results are confirmed by another set of quaternions and spin-rate vectors independently estimated from magnetometer and horizon crossing indicator data. Attitude precision on the order of several degrees has been demonstrated.
29 Jul 2014
TL;DR: In this article, the authors investigated the combustion behavior of paraffin-based hybrid rocket fuels with gaseous oxygen (GOX) as oxidizer and found an exponential relation between the liquid fuel layer viscosity and regression rate, which can be used to predict the regression rate of new liquefying fuels.
Abstract: This paper summarizes the investigations on the combustion behavior of paraffin-based
hybrid rocket fuels with gaseous oxygen (GOX) as oxidizer. Combined experimental activities have been done at the DLR Institute of Space Propulsion in Lampoldshausen and at
the Space Propulsion Laboratory (SPLab) of Politecnico di Milano. Regression rate tests have been done in a 2D radial micro burner at the DLR and at the SPLab. Fuel samples
have been characterized by viscosity measurements, tensile tests and differential scanning calorimeter (DSC). Tensile tests shows significant improvement in maximum stress and elongation when polymers in low concentration are added to the paraffin samples. The values of the liquid fuel viscosities differ signifcantly between the fuels. This affects the droplet entrainment process during combustion and also the regression rates of the fuels. Entrainment and regression rate increase for decreasing fuel liquid layer viscosity. An exponential relation has been found between the liquid fuel layer viscosity and the Regression rate, which can be used to predict the regression rate of new liquefying fuels by measuring their viscosity.
TL;DR: In this paper, an opposed flow burner is examined as an instrument to screen and characterize fuel before full-scale hybrid rocket testing, and an analysis to relate the results to hybrid rocket applications is developed.
Abstract: An opposed flow burner is examined as an instrument to screen and characterize fuel before full-scale hybrid rocket testing. This device requires small amounts (∼10 g) of solid fuels, and it can save time and material in early phases of fuel characterization. Although impinging jet configurations have been investigated in the past, the full range of operation of these systems in terms of hybrid rocket motor flowfield conditions has not been fully explored. The regression rate, flame structure, and flame temperature in an opposed burner configuration is investigated, and an analysis to relate the results to hybrid rocket applications is developed. Hydroxyl-terminated polybutadiene, dicyclopentadiene, and paraffin are investigated via an opposed flow burner over an oxidizer mass flux range of 4 to 25 kg/s/m2. Results show solid-fuel regression rate sensitivity to laminar and turbulent flow regimes. Aluminized hydroxyl-terminated polybutadiene regresses ∼34% slower than neat fuel in the opposed flow burner...