Showing papers on "Liquid-propellant rocket published in 2017"
TL;DR: In this paper, a 6-step Jones-Lindstedt mechanism with an eddy-dissipation concept model was used for turbulent kinetic reaction in bipropellant applications.
57 citations
01 Jan 2017
TL;DR: The first € price and the £ and $ price are net prices, subject to local VAT as discussed by the authors, and the first £ price is net price subject to £ and £ price.
Abstract: The first € price and the £ and $ price are net prices, subject to local VAT. Prices indicated with * include VAT for books; the €(D) includes 7% for Germany, the €(A) includes 10% for Austria. Prices indicated with ** include VAT for electronic products; 19% for Germany, 20% for Austria. All prices exclusive of carriage charges. Prices and other details are subject to change without notice. All errors and omissions excepted. L.T. De Luca, T. Shimada, V.P. Sinditskii, M. Calabro (Eds.) Chemical Rocket Propulsion
52 citations
10 Jul 2017
TL;DR: In this article, the authors describe the design, development, manufacturing and testing of these numerous combustion chambers, and the associated lessons learned throughout their design and development processes, as well as their corresponding hot-fire tests.
Abstract: NASA and industry partners are working towards fabrication process development to reduce costs and schedules associated with manufacturing liquid rocket engine components with the goal of reducing overall mission costs. One such technique being evaluated is powder-bed fusion or selective laser melting (SLM), commonly referred to as additive manufacturing (AM). The NASA Low Cost Upper Stage Propulsion (LCUSP) program was designed to develop processes and material characterization for GRCop-84 (a NASA Glenn Research Center-developed copper, chrome, niobium alloy) commensurate with powder bed AM, evaluate bimetallic deposition, and complete testing of a full scale combustion chamber. As part of this development, the process has been transferred to industry partners to enable a long-term supply chain of monolithic copper combustion chambers. To advance the processes further and allow for optimization with multiple materials, NASA is also investigating the feasibility of bimetallic AM chambers. In addition to the LCUSP program, NASA’s Marshall Space Flight Center (MSFC) has completed a series of development programs and hot-fire tests to demonstrate SLM GRCop-84 and other AM techniques. MSFC’s efforts include a 4,000 pounds-force thrust liquid oxygen/methane (LOX/CH4) combustion chamber. Small thrust chambers for 1,200 pounds-force LOX/hydrogen (H2) applications have also been designed and fabricated with SLM GRCop-84. Similar chambers have also completed development with an Inconel 625 jacket bonded to the GRCop-84 material, evaluating direct metal deposition (DMD) laser- and arc-based techniques. The same technologies for these lower thrust applications are being applied to 25,000-35,000 pounds-force main combustion chamber (MCC) designs. This paper describes the design, development, manufacturing and testing of these numerous combustion chambers, and the associated lessons learned throughout their design and development processes.
49 citations
TL;DR: In this paper, a pintle injector was used for developing throttleable rocket engines, where variable-area injectors are suitable choices for developing a rocket engine because it is difficult to efficiently control thrust when fixed-area injectionors are used.
Abstract: Variable-area injectors are suitable choices for developing throttleable rocket engines because it is difficult to efficiently control thrust when fixed-area injectors are used. A pintle injector i...
41 citations
TL;DR: In this paper, the relative weight of convective and radiative wall heat transfer in liquid rocket engine thrust chambers is estimated by means of dedicated computational fluid dynamics tools, and the relative heat transfer of a single-stage rocket engine is investigated.
Abstract: The relative weight of convective and radiative wall heat transfer in liquid rocket engine thrust chambers is estimated by means of dedicated computational fluid dynamics tools. In particular, alth...
30 citations
TL;DR: In this article, a review of specific features of experimental development of liquid rocket engines (LRE) is presented, including the 11D56, 11D57, RD0120, KVD1, and a number of propulsion units and power plants.
22 citations
TL;DR: In this article, the performance of the noncontact mechanical seal of a high-speed turbopump in a liquid rocket engine operates under very harsh conditions (such as rapid start-up, cryogenic, high speed, high pressure and low-viscosity sealing fluid).
Abstract: The non-contact mechanical seal of a high-speed turbopump in a liquid rocket engine operates under very harsh conditions (such as rapid start-up, cryogenic, high-speed, high-pressure and low-viscosity sealing fluid). The performance of the seal is very different to the performance under normal running conditions. In this paper, for the sake of safety, an experiment is carried out with liquid nitrogen as the sealing fluid. The experimental results with liquid nitrogen are expected to provide an equivalent seal performance as would be experienced with liquid oxygen and liquid hydrogen rocket engine. The main performance parameters, including face temperatures, leakage, face friction force, and friction coefficients, are measured in the speed-up, stable, and speed-down stages. The results show that in the speed-up stage, with rapidly increasing speed, the local face temperature rises dramatically to even higher than the vaporization temperature of liquid nitrogen, and a two-phase flow phenomenon occurs. In the start-up and stable stages, the friction coefficients are 0.25 and 0.13, respectively. After the test, it was found that the wear thickness of the rotor was 0.2 mm, and serious point corrosion appeared on the surface of the stator.
20 citations
10 Jul 2017
TL;DR: In this article, the authors discuss the initial material development, design and fabrication of the subscale carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work.
Abstract: Upper stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures, increasing exhaust velocities. Due to the large size of such nozzles, and the related engine performance requirements, carbon-carbon (C-C) composite nozzle extensions are being considered to reduce weight impacts. Currently, the state-of-the-art is represented by the metallic and foreign composite nozzle extensions limited to approximately 2000 degrees F. used on the Atlas V, Delta IV, Falcon 9, and Ariane 5 launch vehicles. NASA and industry partners are working towards advancing the domestic supply chain for C-C composite nozzle extensions. These development efforts are primarily being conducted through the NASA Small Business Innovation Research (SBIR) program in addition to other low level internal research efforts. This has allowed for the initial material development and characterization, subscale hardware fabrication, and completion of hot-fire testing in relevant environments. NASA and industry partners have designed, fabricated and hot-fire tested several subscale domestically produced C-C extensions to advance the material and coatings fabrication technology for use with a variety of liquid rocket and scramjet engines. Testing at NASA's Marshall Space Flight Center (MSFC) evaluated heritage and state-of-the-art C-C materials and coatings, demonstrating the initial capabilities of the high temperature materials and their fabrication methods. This paper discusses the initial material development, design and fabrication of the subscale carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work. The follow on work includes the fabrication of ultra-high temperature materials, larger C-C nozzle extensions, material characterization, sub-element testing and hot-fire testing at larger scale.
19 citations
01 Jul 2017
TL;DR: The results show that the multi-Algorithm parallel integrated decision-making judgment model gives very effective and reliable performance relative to the voting method, successfully solving multi-algorithm judgment problems and meeting practical engineering needs.
Abstract: This study reports multi-algorithm parallel integrated decision-making for liquid-propellant rocket engine online health condition monitoring to improve reliability and safety, especially for next-generation reusable engines. Fusing multi-algorithm detection information to judge liquid-propellant rocket engine condition is multi-algorithm parallel integrated decision-making main task, and multi-algorithm judgment problem is its central issue; i.e. how to make a global judgment from judgment results of different fault detection methods. Considering opportune fault detection, adequate rocket engine information exploitation, and reliable condition judging, the multi-algorithm parallel integrated decision-making framework for problem definition is presented along with a multi-algorithm parallel integrated decision-making judgment model. For more reliable, efficient global judgment, a method based on the Bayes’ risk function integrating multi-algorithm prior information is adopted. The proposed approach is val...
17 citations
TL;DR: In this article, thermal fatigue panels represent small actively cooled sections of the hot-gas wall of a regeneratively cooled liquid rocket engine, and they are combined with cyclic laser heating.
Abstract: Thermomechanical fatigue panels represent small actively cooled sections of the hot-gas wall of a regeneratively cooled liquid rocket engine. In combination with cyclic laser heating, the panels ar...
16 citations
TL;DR: In this paper, the influence of tip clearance issues in an axial turbine installed to operate as oxidizer booster in the Space Shuttle Main Engine (SSME) was evaluated numerically.
TL;DR: In this paper, a carbon fiber reinforced composite has been adopted to replace the typical heavy metallic closeout structure of a regeneratively cooled thrust chamber of a liquid rocket engine, which provides hoop strength for withstanding the fuel/coolant pressure in the cooling channels.
01 Jan 2017
TL;DR: In this paper, the status of research and development which have been carried out to realize a liquefied natural gas (LNG) rocket engine with higher performance in Japan is reported.
Abstract: This chapter reports the status of research and development which have been carried out to realize a liquefied natural gas (LNG) rocket engine with higher performance in Japan. As a fuel of rocket engine, LNG has better characteristics, i.e., longer storage, lower cost, and nontoxic; hence, LNG rocket engines have been investigated among many countries. However, LNG rocket engines have never been used for actual flight application in Japan, even in the world. The reason is that the performance and characteristics of the current LNG rocket engines do not have enough advantages compared with other liquid rocket engines. For example, if it would apply to a future reusable liquid rocket booster, the specific impulse (Is) of LNG rocket engine should be higher than 360 s in order to get more advantages than other liquid rocket engines. The Is of the current LNG rocket engines in Japan is between 310 and 350 s, falling short of the target level of 360 s. Therefore, continuous research and development have been conducted for the purpose of extending the advantages and promoting the practical use of LNG rocket engines in Japan.
10 Jul 2017
TL;DR: The purpose of the present study is to investigate and compare the major differences of the plume gas flow behaviors numerically between the small monopropellant and bipropellants thrusters and is expected to provide useful information on selecting the appropriate propulsion system.
Abstract: In general, a space propulsion system has a crucial role in the normal mission operations of a spacecraft. Depending on the types and number of propellants, a monopropellant and a bipropellant thrusters are mostly utilized for low thrust liquid rocket engines. As the plume gas flow exhausted from these small thrusters expands freely in a vacuum space environment along all directions, adverse effects of the plume impingement onto the spacecraft surfaces can dramatically reduce the function and performance of a spacecraft. Thus, the purpose of the present study is to investigate and compare the major differences of the plume gas flow behaviors numerically between the small monopropellant and bipropellant thrusters. To ensure efficient numerical calculations, the whole physical domain was divided into three different subdomains depending on the flow conditions, and then the appropriate numerical methods were combined and applied for each subdomain sequentially. With the present analysis results, the plume gas behaviors including the density, the overall temperature and the separation of the chemical species are compared and discussed between the monopropellant and the bipropellant thrusters. Consequently, the present results are expected to provide useful information on selecting the appropriate propulsion system, which can be very helpful for actual engineers practically during the design process.
10 Jul 2017
TL;DR: In this paper, the HEROS 3 was launched from the ESRANGE Space Center to an apogee altitude of 32,300m (106,000 ft) and set a new European student and amateur rocketry and a world altitude record for hybrid sounding rocket built by students.
Abstract: The inherent safety of hybrid rocket propulsion offers some unique advantages com-
pared to solid and liquid propellant rocket engines. This makes it especially attractive for
space tourism, Micro-launcher and hands-on experiments in the education of students. On
November 8th, 2016 at 10:30 a.m. the hybrid sounding rocket HEROS 3 was launched from
the ESRANGE Space Center to an apogee altitude of 32,300m (106,000 ft). This set a new
altitude record for European student and amateur rocketry and a world altitude record for
hybrid rockets built by students. The 7.5m long rocket was using Nitrous Oxide (N2O) and
a Paraffin-based fuel to produce 10,000N of thrust. The dry mass of the rocket was only
75 kg thanks to a carbon fibre structure for the most part. The rocket performed the record
breaking flight at perfect weather and visibility conditions, reaching a maximum airspeed
of 720 m/s and Mach 2.3. The rocket performed a soft landing with two parachutes and
can be reused. Flight data and engine performance data are published and analyzed. The
flight data shows excellent stability of the rocket. Engine performance data proves very
high efficiency and stable combustion as in the ground tests. The subsystem design and
verification before the launch is reported. Engine and flight trajectory simulations show
very good agreements with the flight data. Furthermore, the overall project, the rocket
design, the subsystems as well as the launch campaign are presented here in detail.
TL;DR: In this paper, the analysis of technical requirements (TR) for the development of modern space launch vehicles (LV) with main liquid rocket engines (LRE) is fulfilled in relation to the anthropogenic impact decreasing.
TL;DR: In this paper, an improved laser cycling experiment was designed to reproduce the loads in the rocket combustion chamber in small-scale experiments in full-scale rocket combustion chambers, and the results showed that the heat flux through the coating as well as the shape of the substrate have a great influence on whether the mechanical loads are tensile or compressive.
Abstract: To protect the copper wall of the combustion chamber of regeneratively cooled liquid rocket engines, a coating system may be applied Due to the high cooling heat flux in the combustion chamber wall, a large temperature difference between the hot coating and the cold substrate leads to in-plane stresses as a result of different thermal expansion On the hot side of the wall, these stresses are compressive in the heating phase and become tensile after cooling if the compressive stresses relax at high temperatures To investigate the influence of these loads and to test possible coating systems for the use in rocket engines, laser-cycling experiments were carried out previously The coatings failed mainly by buckling in these experiments, whereas vertical cracks were observed in coatings in full scale rocket combustion chambers The present work elucidates this different behaviour with FEM simulations of the laser cycling experiment and of a coated rocket combustion chamber It was found that the heat flux through the coating as well as the shape of the substrate have a great influence on whether the mechanical loads are tensile or compressive Based on these results, an improved laser cycling experiment was designed to reproduce the loads in the rocket combustion chamber in small scale experiments
TL;DR: In this paper, the analysis of a transverse combustion instability in a reduced-scale rocket engine was conducted on a time-resolved database of three-dimensional fields obtained via large-eddy simulation.
Abstract: This work presents the analysis of a transverse combustion instability in a reduced-scale rocket engine. The study is conducted on a time-resolved database of three-dimensional fields obtained via large-eddy simulation. The physical mechanisms involved in the response of the coaxial hydrogen/oxygen flames are discussed through the analysis of the Rayleigh term in the disturbance-energy equation. The interaction between acoustics and vorticity, also explicit in the disturbance-energy balance, is shown to be the main damping mechanism for this instability. The relative contributions of Rayleigh and damping terms, depending on the position of the flame
with respect to the acoustic field, are discussed. The results give new insight into the phenomenology of transverse combustion instabilities. Finally, the applicability of spectral analysis on the nonlinear Rayleigh and dissipation terms is discussed.
24 Oct 2017
TL;DR: In this article, the authors extended the conventional uncoupled spray model for impinging injectors by considering the coupling of the jet impingement process and the ambient gas field.
30 Sep 2017
TL;DR: In this paper, an improved liquid rocket engine cycle with regenerative cooling of the thrust chamber by oxygen and subsequent use of this oxygen for driving one or two oxygen pumps was proposed and analyzed via comparison with existing staged combustion and gas-generator cycles.
10 Feb 2017
TL;DR: In this article, the authors focus on the computation methods of exhaust plume's flow field, spectral parameters and radiation transfer equation and compare, analyze and conclude the results of these methods.
Abstract: At present, there are various methods to compute the infrared radiation characteristics of exhaust plume of the liquid rocket engine. Though they are different in computational complexity. Their ideas and methods are alike. This paper focuses on the computation methods of exhaust plume’s flow field, spectral parameters and radiation transfer equation. Comparison, analysis and conclusion of these methods are presented. Furthermore, existing problems and improvements of them are proposed as well.
Patent•
11 Jul 2017
TL;DR: In this article, an anti-backfire injector for a single-component rocket engine was proposed, where a micro-pore structure of the injector was proposed to prevent flame spreading up.
Abstract: The invention provides an anti-backfire injecting device for a single-component engine. The anti-backfire injecting device comprises a injecting core body, a cooling supporting ring, an injector and an igniter, wherein a supply runner and a cooling cavity are arranged on the injecting core body; the cooling supporting ring is arranged in the cooling cavity; the injector is mounted on the injecting core body through the cooling supporting ring; and a propellant sequentially flows through the supply runner, the cooling cavity and the injector. The anti-backfire injecting device has the advantages that: the structure is simple, the process is mature and heat is quickly dissipated through a micro-pore structure of the injector, so that flame is prevented from spreading up, and an anti-backfire function and an efficient combustion function are realized; the anti-backfire problem of the propellant supplied by a liquid phase and a gas-liquid phase is solved; the cooling cavity realizes contact heat exchange of the propellant and an ignition plug, so that heat of the ignition plug is absorbed, and the cooling function of the ignition plug is realized; and the anti-backfire injecting device is suitable for a single-component liquid rocket engine and a combustion device, which are liable to generate backfire.
TL;DR: It is concluded that a careful understanding on the plume impingement effects depending on the chemical characteristics of different propellants are necessary for the efficient design of the spacecraft.
Abstract: A space propulsion system is important for the normal mission operations of a spacecraft by adjusting its attitude and maneuver. Generally, a mono- and a bipropellant thruster have been mainly used for low thrust liquid rocket engines. But as the plume gas expelled from these small thrusters diffuses freely in a vacuum space along all directions, unwanted effects due to the plume collision onto the spacecraft surfaces can dramatically cause a deterioration of the function and performance of a spacecraft. Thus, aim of the present study is to investigate and compare the major differences of the plume gas impingement effects quantitatively between the small mono- and bipropellant thrusters using the computational fluid dynamics (CFD). For an efficiency of the numerical calculations, the whole calculation domain is divided into two different flow regimes depending on the flow characteristics, and then Navier-Stokes equations and parallelized Direct Simulation Monte Carlo (DSMC) method are adopted for each flow regime. From the present analysis, thermal and mass influences of the plume gas impingements on the spacecraft were analyzed for the mono- and the bipropellant thrusters. As a result, it is concluded that a careful understanding on the plume impingement effects depending on the chemical characteristics of different propellants are necessary for the efficient design of the spacecraft.
09 Jan 2017
TL;DR: In this paper, the reacting flow from a single gas-centered, swirl-coaxial injector was studied in an optically accessible, high-pressure chamber, with and without high-frequency acoustic perturbations.
Abstract: : The reacting flow from a single gas-centered, swirl-coaxial injector was studied in an optically accessible, high-pressure chamber, with and without high-frequency acoustic perturbations. The gas-centered, swirl-coaxial injector employed liquid rocket engine relevant propellants of gaseous oxygen and RP-2. The reacting flow field behavior at an operating chamber pressure of 3.2 MPa and varying momentum flux ratios were investigated. High-speed shadowgraph images along with OH* and CH* chemiluminescence images were taken to capture the liquid fuel film, droplets, and flame response under acoustic excitation. For the acoustic forcing studies, low amplitude transverse standing waves typically below 5 of the chamber pressure were generated to simulate transverse combustion instabilities. Proper orthogonal decomposition and dynamic mode decomposition were performed on the high-speed shadowgraph and chemiluminescence images to detect the flame response to acoustic forcing, to which in-plane flapping motion was observed for acoustic forcing and rotating soot clouds were a large structures associated with the reacting flow field.
01 Jan 2017
Patent•
05 Apr 2017
TL;DR: In this article, an electric drive propellant feeding system liquid-propellant rocket engine is described, which consists of an oxidizing agent conveying system, a fuel conveying and a thrust chamber.
Abstract: The invention discloses an electric drive propellant feeding system liquid-propellant rocket engine which comprises an oxidizing agent conveying system, a fuel conveying system and a thrust chamber. The oxidizing agent conveying system and the fuel conveying system are independently communicated with the thrust chamber, so that an oxidizing agent and fuel are correspondingly fed into the thrust chamber, are mixed in the thrust chamber and burn in the thrust chamber, and reaction thrust is generated. The oxidizing agent conveying system comprises an oxidizing agent battery, an oxidizing agent motor and an oxidizing agent pump which are sequentially connected, and the oxidizing agent pump is used for being connected with an oxidizing agent storage tank and a thrust chamber pipe. The electric drive propellant feeding system liquid-propellant rocket engine is simple in structure, high in reliability and miniaturized.