Browse Topic: Liquid propellant rocket engines
This document defines and illustrates the process for determination of uncertainty of turbofan and turbojet engine in-flight thrust and other measured in-flight performance parameters. The reasons for requiring this information, as specified in the E-33 Charter, are: determination of high confidence aircraft drag; problem rectification if performance is low; interpolation of measured thrust and aircraft drag over a range of flight conditions by validation and development of high confidence analytical methods; establishment of a baseline for future engine modifications. This document describes systematic and random measurement uncertainties and methods for propagating the uncertainties to the more complicated parameter, in-flight thrust. Methods for combining the uncertainties to obtain given confidence levels are also addressed. Although the primary focus of the document is in-flight thrust, the statistical methods described are applicable to any measurement process. The E-33 Committee
The primary objective of any test program is to maximize the probability, within programmatic constraints, that the flight design will function properly and successfully when used in actual service for the intended application. Flight risks are mitigated via prudent and effective analysis and testing. While analysis can sometimes be used in place of test, proper analytical techniques utilize test data as the basis for model correlations. The combination of analysis and test verification is used for both qualification of the LRE design as well as workmanship verification of each LRE flight unit
An orifice element is commonly used in liquid rocket engine test facilities either as a flow metering device, a damper for acoustic resonance, or to provide a large reduction in pressure over a very small distance in the piping system. The orifice as a device is largely effective in stepping down pressure; however, it is also susceptible to a wake-vortex instability that generates pressure fluctuations that propagate downstream and interact with other elements of the test facility, resulting in structural vibrations. Exacerbating the situation in cryogenic test facilities is the possibility of the formation of vapor clouds when the pressure in the wake falls below the vapor pressure, leading to cavitation. Cavitation has the potential for highamplitude fluctuations that can cause catastrophic damage to a facility
Many pump vaned diffuser designs are based on existing airfoil designs, with little attention given to the vane leading edge. There is a need for a vaned diffuser leading edge that helps resist flow separation and the resultant poor diffuser pressure recovery. Diffusers in pumps are often working with an incompressible fluid that makes potential flow methodologies — which have incompressibility as a boundary condition — attractive. The potential flow-based free-streamline analysis methods have been known to improve the aerodynamics of varied components at high incidence angles, such as diffusers, jet engine nacelles, and liquid rocket engine turbopump inducers
Liquid rocket engine injectors can be extremely expensive to manufacture and hard to iterate to achieve high performance. Internal sealing points can also be the source of reliability issues. The technology disclosed here covers the application of a 3D additive manufacturing (AM) process to produce a functional aluminum injector for liquid propellant rocket engines, along with injector and overall engine design features that optimize the application of such processes to improve performance, reliability, and affordability relative to components produced using standard machining processes and designs. Aluminum was used for the injector instead of higher- temperature metals like stainless steel because its thermal conductance properties provide more opportunity to leverage the cooling potential of liquid oxygen and other cryogenic propellants
This information report presents a preliminary discussion of liquid propellant gas generation (LPGG) systems. A LPGG system, as used herein, is defined as a system which stores a liquid propellant and, on command, discharges and converts the liquid propellant to a gas. The LPGG system can interface with a gas-to-mechanical energy conversion device to make up an auxiliary power system. Figure 1 shows a block diagram of LPGG system components which include a propellant tank, propellant expulsion system, propellant control and a decomposition (or combustion) chamber. The purpose of this report is to provide general information on the variety of components and system arrangements which can be considered in LPGG design, summarize advantages and disadvantages of various approaches and provide basic sizing methods suitable for initial tradeoff purposes
A document describes the low-cost manufacturing of C103 niobium alloy combustion chambers, and the use of a high-temperature, oxidation-resistant coating that is superior to the standard silicide coating. The manufacturing process involved low-temperature spray deposition of C103 on removable plastic mandrels produced by rapid prototyping. Thin, vapor-deposited platinumindium coatings were shown to substantially improve oxidation resistance relative to the standard silicide coating
The primary objective of this document is to describe the systematic and random measurement uncertainties which may be expected when testing gas turbine engines in a range of different test facilities. The documentation covers a "traditional" method for estimating pretest uncertainties and a "new" method for computing and comparing posttest uncertainties. To determine these posttest uncertainties, data generated during the AGARD Uniform Engine Test Program (UETP) were analyzed and compared to the pretest estimates. The proposed procedure provides a mechanism for determining the expected accuracy of test results obtained from facilities which were not previously cross calibrated. Furthermore, the method can be used to assist in making cost-effective management decisions on the level of validation/cross calibration necessary when bringing a test facility on line. This document is also intended to act as a guide for improving uncertainty analyses in a broad spectrum of related industries
The liquid rocket engine stability prediction software (LCI) predicts combustion stability of systems using LOX-LH2 propellants. Both longitudinal and transverse mode stability characteristics are calculated. This software has the unique feature of being able to predict system limit amplitude
This information report presents a preliminary discussion of liquid propellant gas generation (LPGG) systems. A LPGG system, as used herein, is defined as a system which stores a liquid propellant and, on command, discharges and converts the liquid propellant to a gas. The LPGG system can interface with a gas-to-mechanical energy conversion device to make up an auxiliary power system. Figure 1 shows a block diagram of LPGG system components which include a propellant tank, propellant expulsion system, propellant control and a decomposition (or combustion) chamber. The purpose of this report is to provide general information on the variety of components and system arrangements which can be considered in LPGG design, summarize advantages and disadvantages of various approaches and provide basic sizing methods suitable for initial tradeoff purposes
This document defines and illustrates the process for determination of uncertainty of turbofan and turbojet engine in-flight thrust and other measured in-flight performance parameters. The reasons for requiring this information, as specified in the E-33 Charter, are: - determination of high confidence aircraft drag - problem rectification if performance is low - interpolation of measured thrust and aircraft drag over a range of flight conditions by validation and development of high confidence analytical methods - establishment of a baseline for future engine modifications This document describes systematic and random measurement uncertainties and methods for propagating the uncertainties to the more complicated parameter, in-flight thrust. Methods for combining the uncertainties to obtain given confidence levels are also addressed. Although the primary focus of the document is in-flight thrust, the statistical methods described are applicable to any measurement process. The E-33
The primary objective of this document is to describe the systematic and random measurement uncertainties which may be expected when testing gas turbine engines in a range of different test facilities. The documentation covers a "traditional" method for estimating pretest uncertainties and a "new" method for computing and comparing posttest uncertainties. To determine these posttest uncertainties, data generated during the AGARD Uniform Engine Test Program (UETP) were analyzed and compared to the pretest estimates. The proposed procedure provides a mechanism for determining the expected accuracy of test results obtained from facilities which were not previously cross calibrated. Furthermore, the method can be used to assist in making cost-effective management decisions on the level of validation/cross calibration necessary when bringing a test facility on line. This document is also intended to act as a guide for improving uncertainty analyses in a broad spectrum of related industries
A comprehensive mathematical model of mass diffusion has been developed for binary fluids at high pressures, including critical and supercritical pressures. Heretofore, diverse expressions, valid for limited parameter ranges, have been used to correlate high-pressure binary mass-diffusion-coefficient data. This model will likely be especially useful in the computational simulation and analysis of combustion phenomena in diesel engines, gas turbines, and liquid rocket engines, wherein mass diffusion at high pressure plays a major role
This paper presents a study of area production in mixing layers undergoing transition to turbulence. These layers evolve from the mixing of two initially segregated counterflowing streams under supercritical conditions. The study may contribute to development of means to control area production in order to increase disintegration of fluids and enhance combustion in diesel, gas turbine, and liquid rocket engines. As used here, “area production” signifies the fractional rate of change of surface area oriented perpendicular to the mass-fraction gradient in a mixing layer. In the study, a database of transitional states obtained from direct numerical simulations of temporal three-dimensional supercritical mixing layers for heptane/nitrogen and oxygen/hydrogen systems was analyzed. A few of the many conclusions drawn from the analysis are that area production is determined more by strain than by compressibility; area is produced by strain and convective effects; area is destroyed by species
A computational fluid dynamics (CFD) code has been developed to enable simulation of spray combustion near the fuel injectors in a liquid-fueled rocket engine. This code reflects the three-dimensional (3D), multiphase nature of the flow field in a rocket engine and is capable of modeling even a flow field as complex as one that results from the use of impingement injectors. Unlike prior spray-combustion codes that emphasize physical constraints at the expense of geometric ones, this code implements a compromise between physical and geometric constraints in order to enable analysis and comparison of the performances of alternative engine designs that involve different injector geometries. In particular, this code was constructed to enable prediction of the interactive effects of injector-element impingement angles and impingement points, momenta of individual orifice flows, and the resulting combusting flow
An imaging method to detect flaws in composite pressure vessels used in the aerospace industry has been developed. Solid-rocket-motor casings and fuel or oxidizer tanks for liquid rocket motors can now be evaluated with the endoscopic shearography inspection device
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