Browse Topic: Liquid propellant rocket engines

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A One-Piece Liquid Rocket Thrust Chamber AssemblyTBMG-537129/1/2025
NASA is developing a lightweight one-piece regeneratively cooled thrust chamber assembly (TCA) for liquid rocket engines. Liquid rocket engines create thrust through the expansion of combusted propellants within the TCA. Standard manufacturing of TCAs involves individually building the injector, main combustion chamber and nozzle, and then bolting or welding the components together at the joints. However, potential seal failures in these complex joints can cause catastrophic explosions, as in the tragedy of the Space Shuttle Challenger.
Design and Numerical Analysis of Double-Base Swirl Injector for Ethanol/Hydrogen-Peroxide Based Liquid Propellant Rocket Engine2023-01-51002/23/2024
This paper presents a study of numerical cold flow analysis of double-base swirl injector design using Ansys Fluent. The study focuses on the design validation and development of double-base liquid-liquid swirl injector for Ethanol(Fuel) and Hydrogen Peroxide(Oxidizer) based liquid propellant rocket engine. The green propellant contains 80% Ethanol (C2H5OH) as fuel and 60% Hydrogen Peroxide (H2O2) as oxidizer. A comprehensive data, obtained from NASA CEARun code, of performance parameters and carbon monoxide and carbon dioxide emission of most commonly used propellant is compared with ethanol and hydrogen-peroxide based propellant is presented for reference. Secondly, the paper presents the theoretical design model of Swirl Injector, and numerical cold flow study of swirl injector model. For this the 3D models of fuel and oxidizer swirl nozzles are designed separately as per the theoretical design parameters. Poly-hexacore type fluent meshing is used to generate valid mesh. A 3-D, Steady State, Pressure based, SST k-omega Turbulence model is used to carry out the cold-flow simulations. The CFD simulations are carried out separately for oxidizer and fuel nozzles and finally combined model is used for final analysis. Also, this paper presents the study of effects of varying contraction angle (α = 50 °, 55 °, 60°) of the swirl chamber and outlet orifice diameter (do = 3.4, 3.6. 3.8 mm) in swirl injector(oxidizer nozzle) using 9 different models. It was found that increasing or decreasing the contraction angle of the swirl chamber results in narrower and wider spray cone angle and also increment and decrement in mass flow rate of the fluid respectively. The final result of this comparative study concluded that nozzle with α = 55 ° and Do = 3.4mm gave the best result out of the 9 different geometric parameters. The theoretical design mass flow rate of fuel and oxidizer is obtained and validated numerically at design pressure drop value of 1.5 bar for both the fuel and the oxidizer nozzle. The post-processing analysis results are presented in the form of contours, streamlines and comparative graphs. The theoretical design model of the swirl injector is validated successfully through numerical analysis of the swirl injector model. Further, this swirl injector model can be developed to carryout experimental tests and validation.
Kumar, G. Dinesh, Agarwal, Abishek Garg
Uncertainty of In-Flight Thrust DeterminationAIR1678B (Current)3/23/2023
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 has endeavoured to gather industry-wide expertise in in-flight measurement and uncertainty analysis to collect and promulgate recommended practices in the subject disciplines. The Committee is organized into subcommittees to address both the analytical and test methodology for determination of in-flight thrust and also the uncertainty of the determination. This document; Uncertainty of In-flight Thrust Determination, AIR1678, addresses the process for determining the uncertainty of in-flight thrust. A companion document, In-Flight Thrust Determination, AIR1703, addresses the basic methodology for determining in-flight thrust. The Committee, after reviewing recommended changes and clarification in definitions and application of statistical uncertainty items, made small revisions to the original document published in 1985. These changes were incorporated into AIR1678 Rev A. This Revision B has the same Scope as preceding versions. The nomenclature and methodology used herein are now consistent with evolving world and national standards promulgated primarily by ISO and ASME.
E-33 In Flight Propulsion Measurement Committee
Experimental Testing of Spray Characteristics of Swirler Injector without Guide Vanes in Liquid Propellant Rocket Engines2022-28-039310/5/2022
This paper studies the “Experimental testing of spray characteristics of swirler injector without guide vanes in Liquid Propellant Rocket Engines”. The purpose of a fuel injector is to inject and control the flow of the propellants into the combustion chamber. It consists of Tangential ports, a swirls chamber, a converging spin chamber, and a discharge orifice. Kerosene as fuel and liquid oxygen as oxidizer were employed here. To pressurize the propellants, nitrogen gas is used and the pressurized pipeline is controlled with an open/close ball valve and measure the inlet pressure using a pressure gauge. When a propellant comes through the inlet has a tangential velocity and it causes the propellants to swirl inside the swirl chamber at the exit the propellant comes with rotational momentum and forms a liquid film and then the friction between the propellant and air accelerates and disrupts the liquid film and converts into very tiny droplets. The exit mass flow rate, spray cone angle, and spray cone length with different inlet pressures and inlet diameters are measured. A graph is plotted between inlet pressure, and spray cone angle, penetration length. The theoretical mass flow rate is calculated using Bernoulli’s equation, taking the coefficient of the discharge as one and plotted a graph with the given data and comparing it with the experimental data. The accurate atomization increases the combustion efficiency, burns the propellant completely without leaving any unburnt propellant in the combustion chamber, and minimizes combustion instability.
Gangarapu, Yagnateja, ASADALI, KOLAR, Pavuluri, Leelamanideep, Kumar, Dinesh
Aerospace & Defense Technology: August 202222AERP088/1/2022
The Digital Cockpit of the Future How Human Machine Interface Technology Will Impact Avionics Displays The Science of Joining Dissimilar Metals in Aerospace Manufacturing Designing A/D Converters for the James Webb Space Telescope Vibration Testing NASA's X-57 Maxwell Electric Aircraft Hybrid Circulators for mmWave Systems Transforming Battlespace Communications Pulsed Exposure Toxicity Testing: Method Development and Initial Evaluation for Stormwater Compliance Preliminary research into modifying whole effluent toxicity (WET) testing protocols designed for continuous flow discharges as applied to episodic and/or ephemeral discharges such as those associated with storm water runoff. Upgrade Transonic Compressor Test Rig Data Acquisition System New software programming using commercial off-the-shelf software for acquiring data, processing data, displaying results, and generating reports is implemented to improve the performance of a legacy system. Electrokinetics Models for Micro- and Nano-Fluidic Impedance Sensors Microfluidics- and nanofluidics-based impedance sensors play an important role in the identification of toxic industrial chemicals and pathogens in the biodetection and biodefense arena, but their efficient modeling and design continues to be a challenge. Evaluation and Test Requirements for Liquid Rocket Engines This Standard establishes test and evaluation requirements related to the development, qualification (or certification), and acceptance (flight production unit) of liquid propellant rocket engines and associated propulsion systems. GPS Radio Occultation and Ultraviolet Photometry-Colocated (GROUP-C) Early Orbit Testing Results After routine flight, capture, and installation, GROUP-C underwent Early Orbit Testing to verify its performance prior to science operations.
Evaluation and Test Requirements for Liquid Rocket EnginesTBMG-463478/1/2022
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.
Analyses Find Cause of Fatal Problem in Rocket Engine CombustorsTBMG-462878/1/2022
A vital piece of gas engines, combustors — the chambers in which the combustion powering the engine occurs — break down due to fatal high-frequency oscillations during the combustion process. Using advanced time-series analyses based on complex systems, researchers have found what causes them, opening up novel paths to solving the problem.
Cladding and Freeform Deposition for Coolant Channel CloseoutTBMG-398129/1/2021
Low-cost, large-scale liquid rocket engines with regeneratively cooled nozzles will enable reliable and reduced-cost access to space. Coolant contained under high pressure circulates through a bank of channels within the nozzle to properly cool the nozzle walls to withstand high temperatures and prevent failure. It has been a challenge to affordably manufacture and close out the intricate nozzle channels.
Green Electric Monopropellant (GEM)-Fueled Pulsed Plasma ThrusterTBMG-339093/1/2019
NASA required a rocket thruster able to produce a number of pulses at high specific impulse at a relatively low voltage (~300 to 400V). The key problem was that existing propellants for liquid-fueled pulsed plasma thrusters (LPPTs) required high voltages to ablate and accelerate the propellant. The Green Electric Monopropellant (GEM)-fueled Pulsed Plasma Thruster (PPT) is a rocket thruster that is able to produce a number of pulses at high specific impulse at a relatively low voltage (~300 to 400V).
Liquid Rocket Engine: Thrust Chamber Parametric ModelingR-46511/15/2018
The great engineering achievement required to overcome most of the challenges and obstacles that prevented turning rocket design from art into science took place in Europe and the United States between the 1930s and the 1950s. With the vast majority of the engines currently in operation developed in the “pre-computer” age, there are new opportunities to update the design methodologies using technology that can now handle highly complex calculations fast. The space sector with an intense focus on efficiency is driving the need for updating, adapting or replacing the old modeling practices with new tools capable of reducing the volume of resources and the time required to complete simulations and analysis. This book presents an innovative parametric model applicable to the project of some elements of the liquid rocket thrust chamber with the level of detail and accuracy appropriate to the preliminary design phase. It addresses the operating characteristics and dimensioning of some thrust chamber elements through a set of equations and parameters, which include thrust or propellant characteristics. The model degree of sophistication was adjusted to the requirements of the Project Life Cycle Phase B, while also enabling quick analysis of new configurations from changes in initial project parameters.
Nardi Rezende, Rene
Freeform Channel Close-Out for High-Performance Heat ExchangersTBMG-288405/1/2018
Large-scale liquid rocket engines with regeneratively cooled nozzles will enable reliable and reduced-cost access to space. The coolant that circulates through the internal channels within the nozzle helps maintain adequate wall temperatures to prevent failure, and is contained under high pressure. It has been a challenge to affordably manufacture the intricate nozzle channels and close-out of these channels. As such, NASA has leveraged cladding and additive manufacturing technology to build the nozzle liner and outer jacket that closes out the channels within, and contains the high-pressure coolant fluid. The new capability reduces the time to fabricate the nozzle and allows for real-time inspection during the build. It has been demonstrated on a series of different alloys, hot-fire testing is nearly complete, and micrograph examination has verified that the bonds are reliable with little deformation to the coolant channels. The technology provides new opportunities to manufacture nozzles and chambers and reduce costs.
Novel Design of an Orifice Control ElementTBMG-2800612/1/2017
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.
Low Solidity Vaned Diffuser (LSVD) Design for Improvement of Pressure RecoveryTBMG-268525/1/2017
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.
Aluminum Rocket Engine Injector Fabricated Using 3D Additive ManufacturingTBMG-221946/1/2015
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.
Liquid Propellant Gas Generation SystemsAIR1343B (Current)6/17/2013
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-6C4 Power Sources Committee
Advanced Materials and Manufacturing for Low-Cost, High-Performance Liquid Rocket Combustion ChambersTBMG-154991/1/2013
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.
Estimation of Measurement Uncertainty in Engine Tests Based on NATO AGARD Uniform Engine Test ProgramAIR4979A (Current)5/29/2012
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. Measurement uncertainty and measurement system approaches and practices for the UETP are presented in a systematic format. Readers can use these uncertainty approaches and practices for comparison to their own measurement systems. The approach chosen was to analyze three of the UETP test conditions at eight test facilities along with the individual measurement system approaches and practices. The test program provided a wide range of performance measurements and corresponding pretest uncertainty estimates encompassing a variety of test and measurement approaches and practices. Included in the analysis are the basic gas turbine performance measurements of temperature, pressure, airflow, fuel flow, area, speed, thrust, and typical performance parameter functions and associated uncertainty estimates.
E-33 In Flight Propulsion Measurement Committee
Integration of Probabilistic Methods into the Design ProcessAIR5080 (Current)5/8/2012
This document discusses a recommended new approach to integrate probabilistic methodologies with design practices, procedures, and software codes currently being used. In addition to complementing design methods currently in use, this new procedure will permit the designer to quantify the amount of conservatism that exists for a particular design due to the large amount of additional information which is provided to the designer. This additional information will allow the designer to make better decisions when faced with tradeoffs between cost, reliability, performance, and weight. Although the methodologies described herein can be used heavily in the design process, their applicability is much more encompassing. They can be used from product concept to customer delivery.
G-11 Probabilistic Methods and Uncertainty Quantification
Nonlinear Combustion Instability PredictionTBMG-84329/1/2010
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.
Liquid Propellant Gas Generation SystemsAIR1343A (Historical)11/6/2007
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-6C4 Power Sources Committee
Uncertainty of In-Flight Thrust DeterminationAIR1678A (Historical)6/11/2007
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 has endeavored to gather industry-wide expertise in in-flight measurement and uncertainty analysis to collect and promulgate recommended practices in the subject disciplines. The Committee is organized into two subcommittees to address both the analytical and test methodology for determination of in-flight thrust and also the uncertainty of the determination. This document "Uncertainty of In-flight Thrust Determination, AIR1678", addresses the process for determining the uncertainty of in-flight thrust. A companion document, "In-Flight Thrust Determination, AIR1703", addresses the basic methodology for determining in-flight thrust. The Committee, after reviewing recommended changes and clarification in definitions and application of statistical uncertainty items, have made small revisions to the original document published in 1985. These changes will maintain consistency of AIR1678 with evolving world and national standards promulgated primarily by ISO and ASME.
E-33 In Flight Propulsion Measurement Committee
Estimation of Measurement Uncertainty in Engine Tests Based on NATO AGARD Uniform Engine Test ProgramAIR4979 (Historical)11/17/2004
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. Measurement uncertainty and measurement system approaches and practices for the UETP are presented in a systematic format. Readers can use these uncertainty approaches and practices for comparison to their own measurement systems. The approach chosen was to analyze three of the UETP test conditions at eight test facilities along with the individual measurement system approaches and practices. The test program provided a wide range of performance measurements and corresponding pretest uncertainty estimates encompassing a variety of test and measurement approaches and practices. Included in the analysis are the basic gas turbine performance measurements of temperature, pressure, airflow, fuel flow, area, speed, thrust, and typical performance parameter functions and associated uncertainty estimates.
E-33 In Flight Propulsion Measurement Committee
Calculating Mass Diffusion in High-Pressure Binary FluidsTBMG-8155/1/2004
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.
Area Production in Supercritical, Transitional Mixing LayersTBMG-32605/1/2002
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 mass flux, rotational effects, and pressure gradients; area can be either produced or destroyed by pressure gradients; and effects of viscosity on area production are negligible. Effects of departure from perfect-gas and ideal-mixture behavior were found to be important. Smaller-wavelength initial perturbations were found to lead to greater area production: this observation could be a guide to initial development of control of area production.
Software for Simulation of 3D, Three-Phase Combusting FlowTBMG-745310/1/2001
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.
Endoscopic Shearography InspectionTBMG-320144/1/1998
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.
LIQUID ROCKET ENGINE RELIABILITY CERTIFICATIONARP4900 (Current)4/1/1996
Current design and development practices leading to formal liquid rocket engine qualification (USAF) or certification (NASA) will not achieve the specific reliability objectives of future programs. New rocket engine programs are dictating quantified requirements for high reliability in parallel with a cost-constrained procurement environment. These specified reliability levels cannot be validated with the necessary confidence in a timely or cost-effective manner by present methods. Therefore, a new improved process is needed and has been developed. This new reliability certification methodology will be discussed in detail in the five sections that comprise this document. Primary purposes of this report are to: a Define and illustrate this process b Point out its strengths and weaknesses c Provide guidelines for its application on programs which have specified reliability requirements Increased emphasis on rocket engine reliability and cost has prompted the Liquid Rocket Certification Subcommittee (Society of Automotive Engineers for Reliability, Maintainability, and Supportability) to thoroughly examine current methodologies to qualify or certify liquid rocket engine systems. For example, new liquid rocket engine programs, such as the joint NASA/Air Force effort for the National Launch System (NLS) or the Air Force XLR-132 storable propellant upper stage engine, include documented requirements for high levels of reliability. These new requirements exceed those historically demonstrated over the operational life of most current rocket propulsion systems. Certification of reliability was not required for past liquid rocket engines developed for the Air Force or NASA. The importance of demonstrated reliability was low, relative to such requirements as performance, schedule, and cost. Engines were formally qualified or certified by test programs aimed primarily at demonstrating design maturity and operational readiness in terms of performance and durability. In general, relatively little propulsion system testing, as distinguished from engine system testing, was implemented on past flight hardware for launch vehicles. Reliability estimates prior to the first flight of a new engine historically have been based largely upon results from qualification or certification tests which formally declared the engine ready to fly. Many changes typically were made during the engine development period, until the engine was considered mature enough to qualify or certify. The process, therefore, precluded the gathering of test results applicable to reliability assessment during this development phase of a program. As a consequence, predicted reliability levels, at high confidence, prior to the first flight of a new engine have been consistently low. This was due to the small number of engines tested, especially identical units, and the limited number and type of tests performed on each engine during a typical qualification or certification test program. Reliability levels for current operational rocket engines are based upon a combination of ground test experience supplemented by the accumulation of data derived from actual flights. This process typically takes years and hundreds to perhaps thousands of tests to develop a satisfactory level of reliability and confidence for a particular engine system. The Liquid Rocket Certification Subcommittee advocates a new approach to rocket engine reliability certification as a result of reviewing current methods to qualify or certify engines. It is felt that this new approach is an improvement over current qualification/certification methods. The recommended new approach, described in the following sections of this report, involves a judicious combination of analysis and test efforts that begin at an early stage of the design prior to formal certification. This methodology quantifies reliability estimates by focusing upon early identified weak links in the design and system reliability drivers. The recommended approach includes development tests that assist in establishing the necessary information base for probabilistic analyses and engine system certification testing to demonstrate structural, thermal, and dynamic capabilities, as well as the more typical performance and life requirements. The new approach begins with a traditional deterministic preliminary design of the engine. A failure modes and effects analysis and a fault tree analysis are then conducted. At this point, the improved approach departs from typical methodology by screening engine components for criticality. A critical component has one or more critical failure modes. This screening is based upon the accumulated knowledge which impacts the design at this point. Critical components typically are complex in geometry, difficult to analyze, susceptible to catastrophic failure, and sensitive to such things as environments, loads, or material properties. Experience has shown that a majority (about 80 to 90%) of the components of a rocket engine can be classified as noncritical, and their reliability is essentially unity. Therefore, a conventional deterministic design approach is satisfactory for these components. However, probabilistic analysis may be desirable for these noncritical components to realize other benefits such as weight savings. The remaining engine components have a higher probability of failure as well as being engine system critical and require the more intensive probabilistic analysis. A probabilistic analysis recognizes dimensional tolerances, variability in material properties, inadequacies in modeling techniques, load distributions, manufacturing variabilities, and so forth, involved in each critical failure mode. Components that utilize the more intensive probabilistic analysis techniques will yield quantified reliability estimates, while those designed deterministically are assessed only for serviceability. The process is iterative and continuous in nature, whether the component follows the deterministic or probabilistic path, and utilizes the best information available at the time of the analysis. Data deficiencies identified by the probabilistic analysis approach provide guidance for establishing a cost-effective test program during the development phase of the engine program. The final step in the recommended new approach is a formal, hot firing, test of the engine system which simulates, to the maximum extent possible, the complete propulsion system. Tests will be conducted to engine operational limits to validate structural, thermal, and dynamic margins. A careful review of earlier rocket engine certification and re-certification test programs revealed a number of weaknesses in these formal programs. For example, tests were implemented on a very limited number of like engines. Similarly, most tests were conducted at nominal engine operating conditions with little or no testing at or near anticipated flight operational boundaries. Few attempts were made to demonstrate structural, dynamic, or thermal margins. Duration typically was stressed by multiple full-term firings as suggested in MIL-R-5149 (1969). However, margins in duration frequently were compromised by engine rework. Early test programs failed to provide adequate reliability data because of the many shortcomings indicated above. Sections 5 and 6 of this report will be devoted to reliability validation for the application of this new approach on programs that have specified requirements for engine reliability. In summary, an examination of weaknesses in past programs to qualify or certify liquid rocket engines, combined with recent strong emphasis on high engine reliability, has led to this recommendation of a new improved approach for the entire process. It is hoped that this new approach will be adopted by and satisfy the future needs of the military, NASA, and commercial users of liquid rocket engines because of the many advantages that will accrue from this approach. For example, it elevates reliability to a status typical of performance, schedule, and cost. It provides early identification of weak components and mitigates nonbeneficial conservatism due to compounding of margins and factors of safety on some components. The approach also guides cost-effective test programs to validate analytic models, confirm environmental predictions, and define system interactions. It provides continuous quantified estimates of component and engine reliability and validates the required level of reliability prior to commitment to flight. It demonstrates structural, thermal, and dynamic capability to operational limits. Finally, the new approach reduces total costs of development, certification, and flight, at some affordable increase to the initial design costs.
G-11 Probabilistic Methods and Uncertainty Quantification
Health Monitoring of Reusable Rocket Engines9210074/1/1992
An extensive study of liquid propellant rocket engine test and flight history data has produced methods for advanced health monitoring currently under development for application to the Advanced Shuttle, National Aero-Space Plane, and other future propulsion systems. Advanced Health Monitoring is a systematic approach to instrumentation design and development, signal processing, and knowledge-based strategies to infer the operating status of any engine component or system. Properly implemented, Advanced Health Monitoring increases safety and the probability of mission success, significantly reduces turnaround time and maintenance cost, prolongs component life and reduces life-cycle cost.
Ezell, W., Barkhoudarian, S., Cross, G.
The J-2 Liquid Hydrogen Rocket Engine6400401/1/1964
This paper discusses the engine system and component design of the J-2 liquid propellant rocket engine, with emphasis upon the newer features whose development has been dictated by the use of liquid hydrogen as the fuel. A brief description of the development program and of the principal specialized test equipment is given.
Studhalter, W. R.
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