Browse Topic: Liquid propellant rocket engines

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Large Eddy Simulations of Supercritical and Transcritical Jet Flows Using Real Fluid Thermophysical Properties2020-01-11534/14/2020
In order to understand supercritical jet flows further, well resolved large eddy simulations (LES) of a n-dodecane jet mixing with surrounding nitrogen are conducted. A real fluid thermodynamic model is used to account for the fuel compressibility and variable thermophysical properties due to the solubility of ambient gas and liquid jet using the cubic Peng-Robinson equation of state (PR-EOS). A molar averaged homogeneous mixing rule is used to calculate the mixing properties. The thermodynamic model is coupled with a pressure-based solver to simulate multispecies reacting flows. The numerical model is based on a second order accurate method implemented in the open source OpenFOAM-6 software. First, to evaluate the present numerical model for sprays, 1D advection and shock tube benchmark problems at supercritical conditions are shown. Second, a cryogenic nitrogen injection with a jet velocity of 4.9 m/s into a supercritical nitrogen environment at 4 MPa and room temperature is considered to carry out a grid resolution study, and the corresponding results are evaluated against experimental data of Mayer et al. Then, to assess the effects of thermophysical property variations due to mixing of two species, a high-pressure jet of n-dodecane at transcritical and supercritical temperatures at 200 m/s into high-pressure and high-temperature nitrogen environment is studied. Detailed analysis of the species dispersion and mixing are presented for various conditions. The present LES simulations of n-dodecane jets show massive shear forces and high hydrodynamic pressure fluctuations caused by the high-speed jet. The predicted results of flow and thermophysical properties are in close agreement with the available literature data.
Ningegowda, B M, Rahantamialisoa, Faniry, Zembi, Jacopo, Pandal, Adrian, Im, Hong G., Battistoni, Michele
Effects of Engine Speed on the Performance at Extreme Vehicle Driving Conditions2019-36-02971/13/2020
Nowadays, improvements in engine fuel economy and reduction in pollutant emissions have been much discussed. Downsizing and downspeeding are methods widely used in the automotive market, used to increase internal combustion engine efficiency. Fuel consumption is commonly measured through a specific cycle that could differ between countries. Federal Test Procedure (FTP-75) for Latin America and New European Driving Cycle (NEDC) for Europe are some examples. These cycles include normal driving conditions, low vehicle load and mild environmental conditions. The style of gear shifting is one of the main factors affecting fuel efficiency and performance of motor vehicles. At extreme diving conditions, parameters like coolant and air charge temperature could reach limit durability values of the system. In order to avoid knocking and maintain structural reliability, it becomes mandatory to control the engine speed. An increase in engine speed reduces engine load requirements and thus, turbocharger loads. Lower turbocharger requirements reduce ACT and, consequently, reduce knocking. This process allows an advanced ignition timing, enabling more chamber pressure and improving fuel conversion efficiency. However, a higher engine speed also increases friction losses and combustion frequency, lowering the generated torque and tending to decrease engine efficiency. It is interesting to note parameters that affect positively and negatively at the same time, depending on the point of view. The present paper aims to analyze the performance and fuel consumption of a turbocharged engine working under extreme driving conditions, simulating a CVT transmission behavior at various speeds. The evaluated vehicle parameters are: power, coolant temperature, ACT and brake specific fuel consumption (BSFC).
Thomaz, Fabrício, Baeta, José Guilherme Coelho
A Target Cascading Method Using Model Based Simulation in Early Stage of Vehicle Development2019-01-08364/2/2019
In the early stages of vehicle development, it is important for decision makers to understand a feasible constraint region that satisfies all system level requirements. The purpose of this paper is to propose a target cascading method to solve for a feasible design region which satisfies all constraints of the system based on model based simulation. In this method, the feasible design region is explored by using both global optimization methods and active learning techniques. In optimization problems, the inverse problem for understanding feasibility for specific designs is defined and solved. To determine the objective functions of the inverse problem, an index representing the achievement level of constraints from system requirements is introduced. To predict feasible regions in the specific design space, a surrogate model of minimized values of the index is trained by using a kriging model. Training data of the surrogate model is sequentially generated based on the expected improvement function to improve accuracy of feasible regions effectively.
Shintani, Kohei, Abe, Atsuji, Yamamoto, Yasushi
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.
Effect of High Frequency Acoustic Field on Atomization Behavior of Ethanol and Kerosene2017-01-231810/8/2017
Combustion instability often occurs inside the combustion chamber of aero engine. Fuel atomization and evaporation, one of the controlling processes of combustion rate, is an important mechanism of the combustion instability. To tackle combustion instability, it challenges a deep understanding of the underlying mechanism of fuel atomization and evaporation. In this paper, acoustic field was established to simulate the pressure oscillation. Transient spray images of ethanol and kerosene were recorded using high-speed camera. The obtained images were processed by MATLAB to extract and analyze the related data. Spatial fuel atomization characteristics was analytically examined by multi-threshold image method to analyze the effect of the high frequency acoustic field on the fuel break-up and disintegration. The results show that the half spray cone angle on the side with speaker is suppressed by the presence of the imposed acoustic field compared with the case without speaker. Statistically, the half spray angle of kerosene with right speaker under the acoustic frequency of 9 kHz is 18.97% larger than that with left speaker. For ethanol, the difference is 11.90%. Also, it turns out the frequency of the acoustic field influences the spray angle variation during the injection process. There is a decline of the spray cone angle at the early stage of the injection. However, the watersheds for the transformation were not the same for kerosene and ethanol.
Jia, Xiaoxu, Huang, Zhong, Ju, Dehao, Huang, Zhen, Lu, Xing-cai
Numerical Analysis on the Injection and Atomization Characteristics of Diesel Surrogates at Engine Conditions2017-01-230610/8/2017
Recently, the shortage of fossil resources contributes to strict regulations of environmental protection. The research on the high efficiency and low emission of engines becomes an important direction all over the world. Technologies like high injection pressure, high levels of supercharging and higher levels of back pressure have come into application. Increasing the injection pressure and average cylinder pressure results in that parts of the spray can experience transcritical and supercritical regimes. In this paper, we established a surrogate fuel composed of n-Hexadecane, HMN and 1-Metylnaphthalene, to analyze the injection and atomization of diesel surrogate fuel with large eddy simulation (LES) in a cubic calculation region with high temperature and high pressure environment. The injection pressure was fixed to 150MPa, and the 900-K temperature and the 6-MPa pressure represented the ambient condition in constant volume vessel which is supercritical with respect to No.2 diesel. Analyses of No.2 diesel and the surrogate fuel were performed to compare the characteristics of injection and atomization. To illustrate the differences, liquid penetration, jet penetration and spray cone angle of these two fuels were studied and discussed, and the influencing factors like distillation temperature, kinematic viscosity and density were considered as important reasons. The result shows that when the surrogate fuel is injected into the ambient under supercritical conditions, the liquid penetration is slightly longer than that of No.2 diesel. The jet penetration of No.2 diesel is slightly longer than that of the surrogate fuel. Moreover, the spray cone angle of the surrogate fuel is wider than that of No.2 diesel. In summary, the fuels of different thermophysical characteristics may develop and diffuse differently resulting in different characteristics of injection and atomization.
Sun, Xiaochuan, Li, Xiang, Huang, Zhong, Ju, Dehao, Lu, Xing-cai, Han, Dong, Huang, Zhen
Application of Solid State Welding and Superplastic Forming to Aerospace Vehicles2017-01-21489/19/2017
In this paper, manufacturing of launcher structural components with superplastic forming (SPF) and solid state welding technologies is presented with several examples. Some of high strength aerospace alloys, like aluminum, titanium and superalloys, are known to have superplasticity so that complex shapes of aerospace components can be produced with this technology. A combination of superplastic forming and solid state welding processes produces lighter and stiffer components than one manufactured with conventional machining and welding. Solid state welding is an attractive method to weld materials without melting where mechanical properties are important since the welding interface is homogeneous without liquid phases. The results includes aluminum and titanium tanks cryogenic application, a scaled combustion chamber with duplex stainless steel and copper alloy, and lightweight aluminum-copper-lithium cylinder manufactured by superplastic forming and diffusion bonding or friction stir welding.
Lee, Ho-Sung
Combustion Characteristics of Hydrocarbon Droplets Induced by Photoignition of Aluminum NanoparticlesTBMG-273818/1/2017
In the study of combustion characteristics of liquid rocket fuels, it is customary to either study the combustion of liquid fuel droplets or the combustion of fuel sprays. However, the two are closely related to each other, because in a typical rocket combustion chamber, the burning of droplets, droplet clusters, and fuel sprays occur simultaneously.
Combustion Characteristics of Hydrocarbon Droplets Induced by Photoignition of Aluminum Nanoparticles17AERP08_076/1/2017
Test methodology allows analysis of combustion dynamics for subscale rocket injectors under super critical conditions. Air Force Research Laboratory (AFMC), Edwards Air Force Base, California In the study of combustion characteristics of liquid rocket fuels, it is customary to either study the combustion of liquid fuel droplets or the combustion of fuel sprays. However, the two are closely related to each other, because in a typical rocket combustion chamber, the burning of droplets, droplet clusters, and fuel sprays occur simultaneously. The study of droplet burning is mostly focused either on single suspended droplets or free droplets, usually in a free fall. In either case, the conditions regarding the combustion of a single fuel droplet deviates from what is expected in typical liquid rocket engines (LREs). This not only concerns the chamber pressure and the temperature that is usually much lower in a typical droplet study, but also the droplet ignition process and the related transient effects such as temperature ramp up that is much slower than what takes place in LREs. As a result, reported values of combustion parameters such as the burning rate constant and the ignition delay time obtained under typical experimental conditions may not entirely represent the combustion processes in LREs.
Green Monopropellant Secondary Payload Propulsion SystemTBMG-268355/1/2017
Small satellites, launched as secondary payloads, are increasingly being fielded. Advances in liquid rocket propulsion that enhance the on-orbit maneuverability, increase the on-orbit life, and decrease the time between identified need for and deployment of such spacecraft are of great value. Replacing the nearly ubiquitous yet toxic hydrazine propellant with AF-M315E produces higher specific impulse and density specific impulse, resulting in improved overall velocity change capability and increased on-orbit life.
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.
Study of Formulating Conserved Scalar Equations for Turbulent Reactive Flows with General Species Mass- Diffusion Coefficients for Utilization in Flamelet ModelsTBMG-261651/1/2017
The modeling of turbulent reactive flows is a subject of contemporary research. Current turbulent-reaction models cannot account for realistic complexities such as distinct species mass-diffusion coefficients. Under the assumption of a single, constant, mass-diffusion coefficient, a conserved-scalar equation is typically derived in turbulent reactive flows by taking the difference between chemical-species conservation equations having opposite reaction rates (in the sense that the reactant has an opposite reaction rate to the product), thereby creating an equation devoid of reaction terms. Assuming the reaction regions are very thin and are merely contorted by turbulence, chemistry and turbulence can be decoupled, and the evolving statistics of the conserved scalar describe the reaction progression. No such equation has yet been derived for distinct mass-diffusion coefficient cases where the single coefficient is now replaced by a full matrix. Considering that mass diffusion is responsible for reactants approaching at the molecular level and for reaction initiation, this lack of mathematical framework is very disturbing.
NASA's Software Answers Big Questions About Space and EarthTBMG-2606312/1/2016
NASA’s next-generation Space Launch System (SLS) for deep space exploration consists of four RS-25 liquid rocket engines and two five-segment solid rocket boosters (SRBs). At ignition, the SRBs create a significant over-pressure event, known as the Ignition Over-Pressure (IOP) event. The IOP event experienced in the first space shuttle flight (STS-1) in 1981 damaged the space shuttle orbiter and motivated significant design changes to the launch pad at NASA’s Kennedy Space Center. The RS-25 liquid rocket engines aboard the SLS, along with the SRBs, also create a very significant acoustic environment that, if not mitigated, can harm the SLS launch vehicle. An Ignition Over-Pressure/Sound Suppression (IOP/SS) water system was developed to reduce the amplitude of both the IOP event and the acoustic environment caused by the firing of the RS-25 engines and the SRBs.
Uncertainty of In-Flight Thrust DeterminationAIR1678B (Current)10/22/2016
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
Eddy-Current-Minimizing Flow Plug for Use in Flow Conditioning and Flow MeteringTBMG-2551010/1/2016
Innovators at NASA’s Marshall Space Flight Center have developed a suite of prototype fluid plug technologies with an array of capabilities for fluid flow metering, mixing, and conditioning. Each innovation within this suite is based upon a core technology that has no moving parts, is simple to manufacture, and provides high reliability and efficiency. Also, the base fluid plug technology can be modified with very few or no hardware changes to achieve the desired effect or combination of mixing, metering, and conditioning capabilities depending on the application.
Distributed, Fast, Intelligent Infrastructure Health Monitoring SystemTBMG-242674/1/2016
NASA needed an innovative solution to conduct system assessments of frameworks and to characterize subsystems of interest. This tool would be required to determine anomalies; examine their causes (root-cause analysis); make predictive statements (prognostics); provide intelligent health monitoring with incremental knowledge for working with unknown scenarios; and provide maintenance support for a complex collection of systems, subsystems, and elements in rocket engine test platforms.
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.
Advanced Numerical Tools for Design and Analysis of In-Space Valve and Feed SystemsTBMG-205549/1/2014
A high-fidelity numerical simulation software (CRUNCH CFD®) predicts the transient performance of flight valve designs, provides design support by supplementing current empirical rules, and diagnoses system anomalies. Currently, transient analysis of valves is difficult to simulate because of the requirement to dynamically deform the grid due to the valve motion. For complex, transient problems such as engine startup or shutdown that also involve dynamic sealing of fluid flow paths due to valve/solid surface contact, it becomes nearly impossible to deform the grid in an automated fashion.
3-D Manufacturing of Titanium Components Takes OffTBMG-203428/1/2014
With such challenges as base closings, shrinking defense budgets, and sequestration, the worldwide maintenance, repair, and overhaul (MRO) sector is projected to experience a significant decline over the next few years.
Elevated-Temperature, Highly Emissive Coating for Energy Dissipation of Large SurfacesTBMG-199036/1/2014
This coating demonstrates high emittance above 80% or better at broad wavelengths within the infrared spectrum. It has shown to have an extremely stable emittance at lower wavelengths within the infrared (IR) spectrum, where energy dissipation is critical at elevated temperatures. The coating has demonstrated increases in surface texturing, and ultimately an increase in emissivity when exposed to temperatures up to 2,050 °F (≈1,120 °C). It is also stable at continuous run, elevated temperatures, and shows no signs of spalling or erosion.
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
Simulation of Oxygen Disintegration and Mixing With Hydrogen or Helium at Supercritical PressureTBMG-1486810/1/2012
The simulation of high-pressure turbulent flows, where the pressure, p, is larger than the critical value, pc, for the species under consideration, is relevant to a wide array of propulsion systems, e.g. gas turbine, diesel, and liquid rocket engines. Most turbulence models, however, have been developed for atmospheric-p turbulent flows. The difference between atmospheric-p and supercritical-p turbulence is that, in the former situation, the coupling between dynamics and thermodynamics is moderate to negligible, but for the latter it is very significant, and can dominate the flow characteristics. The reason for this stems from the mathematical form of the equation of state (EOS), which is the perfect-gas EOS in the former case, and the real-gas EOS in the latter case.
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
Applications of Probabilistic MethodsAIR5109 (Current)5/8/2012
This SAE Aerospace Information Report (AIR) addresses the following: 1 Captures previous experience and lessons learned in the application of PM. 2 Tabulates public-domain applications, and several representative examples discussed in detail. 3 Notes relative merits and barriers to implementation. The document does not contain technical details of probabilistic methods, benchmarking of specific approaches or legal aspects. These subjects are covered in other AIRs, referenced in Section 2 and prepared by the Probabilistic Methods Committee of the G-11 Reliability, Maintainability, Supportability and Logistics (RMSL) Division of SAE.
G-11 Probabilistic Methods Committee
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
Liquid Oxygen/Liquid Methane Rocket Engine Development2007-01-38769/17/2007
Because of low toxicity and low reactivity, methane and oxygen are identified as “green propellants”, and have the highest performance (as measured in ISP) of any practical rocket propellants, except oxygen/hydrogen. Investigation into green propellant rocket engines identified liquid oxygen (LOX)/methane (CH4) as an excellent candidate for upper stage applications. This paper discusses work in LOX/CH4 engine development carried out by XCOR Aerospace, Inc.
DeLong, Dan, Greason, Jeff, McKee, Khaki Rodway
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
Super-High-Strength Aluminum AlloyTBMG-46586/1/2006
AFRL researchers developed a superhigh- strength aluminum alloy that engineers can use to improve the capability and performance of aerospace components—cryogenic rocket engine components, in particular. They created an aluminum alloy with specific strength and ductility characteristics surpassing those of the alpha titanium alloy currently used in rocket engine turbopumps. The aluminum alloy also demonstrates less sensitivity to hydrogen embrittlement, is lighter weight, and is potentially less costly to manufacture than the titanium alloy.
Theoretical and Experimental Studies of Electrospray for IC Engine2006-01-13884/3/2006
The objective of this work was to investigate the potential of the electrostatic atomization for its application in internal combustion engines. In this paper, a theoretical model for secondary breakup of charged droplets was established. The electric force reduces the surface tension of liquid, whereby atomization is promoted. To improve the diesel droplet atomization remarkably by means of electrostatic charge, the charge-mass ratio should be at least at the order of 10-6C/g. In the interest of the practical application conditions in internal combustion engines, the high-pressure injected electrospray was generated and investigated under various injection pressures and electric conditions. By means of the Photron high-speed camera, the special features of electrospray were observed. The micro-characters including the drop size distribution and the variance of the drop diameter in the spray front area were investigated. The uniformity of the diesel droplet sizes in this area was improved by electrostatic charge significantly.
Li, Liguang, Yu, Shui, Hu, Zongjie
Design of D-B Injectors for Increasing Propulsive Force in Jets2005-01-338810/3/2005
In this work the fundamentals of centrifugal injector calculation is reviewed and new design procedure is proposed for liquid-liquid injectors, based on both theory and experimental results. Then some special conditions related to dual based liquid-liquid injectors are studied and the corresponding results were considered in design manipulation. The behavior of injector in various service conditions, such as mutual effects on twin spray envelopes in dual-base injector is investigated, then the design procedure is presented based on obtained results. Finally the computer program for designing this sort of injector is proposed. Using this program, some injectors were designed and tested. The results were presented in details in this paper.
Ommi, F., Hossinalipour, M., Movahednejad, E.
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.
Update on Area Production in Mixing of Supercritical FluidsTBMG-770011/1/2003
The paper “Turbulence and Area Production in Binary-Species, Super- critical Transitional Mixing Layers” presents a more recent account of the research sum- marized at an earlier stage in “Area Production in Super- critical, Transitional Mixing Layers” (NPO-30425), NASA Tech Briefs, Vol. 26, No. 5 (May 2002) page 79. The focus of this research is on supercritical C7H16/N2 and O2/H2 mixing layers undergoing transitions to turbulence. The C7H16/N2 system serves as a simplified model of hydrocarbon/air systems in gasturbine and diesel engines; the O2/H2 system is representative of liquid rocket engines. One goal of this research is to identify ways of controlling area production to increase disintegration of fluids and enhance combustion in such engines. As used in this research, “area production” signifies the fractional rate of change of surface area oriented perpendicular to the mass-fraction gradient of a mixing layer. In the study, a database of transitional states obtained from direct numerical simulations of the aforementioned mixing layers was analyzed to investigate global layer characteristics, phenomena in regions of high density-gradient magnitude (HDGM), irreversible entropy production and its relationship to the HDGM regions, and mechanisms leading to area production.
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.
Vortex-Fired Liquid-Fuel Rocket Combustion ChambersTBMG-69419/1/2000
Liquid-fuel rocket engines that utilize vortex flow fields to keep combustion-chamber walls cool have been investigated in computational simulations and experiments. In an engine of this type, the vortex flow establishes radial gradients of pressure and density that cause the lower-density hot combustion products to be confined near the central axis of the combustion chamber while cold gases yet to be burned are centrifuged to the combustion-chamber wall. Keeping combustion-chamber walls cool in this manner reduces or eliminates damage by heat and oxidation, thereby (1) eliminating the need for leak-prone cooling passages, (2) extending the operational lifetimes of combustion chambers, and (3) creating an opportunity to fabricate combustion-chamber walls more easily, using relatively inexpensive materials (possibly even lightweight composite materials) instead of expensive alloys.
Refractory Metal-Lined Composites for Lightweight High-Performance Propulsion omponents2000-01-17205/16/2000
Requirements for advanced rocket propulsion systems are becoming increasingly more demanding. The use of high temperature capable materials in such systems, including applications in liquid rocket engines and solid rocket motors, offers potential benefits of increased performance and/or efficiency based on the engine operating cycle. For the ultimate in high temperature capability, refractory metals, ceramics, ceramic matrix composites (CMCs), and carbon/carbon (C/C)composites each provide particularly beneficial attributes, but with selected limitations. Refractory metals are relatively tough and durable, provide impermeable structures, and can be conventionally fabricated, but are relatively dense, leading to heavyweight structures. Monolithic ceramics generally lack desired toughness and durability. CMCs offer substantially improved toughness over their monolithic counterparts and are relatively lightweight, but are permeable and difficult to join to conventional structures. C/C is extremely lightweight, but lacks resistance to the operating environment and is also permeable. Metal-lined, composite-jacketed structures have the potential to combine the beneficial characteristics of metals with the light weight and high temperature resistance of CMCs or C/C. In the current work, prototype refractory metal-lined C/C structures were designed, fabricated, and successfully ground tested in a representative liquid rocket engine environment. No leakage or permeability was exhibited. Successful joining was achieved to the mating structure, and cost-effective fabrication was demonstrated using an “inside-out” processing technique. Projected full-scale component weight was similar to that for conventional materials employed in similar applications. This approach, when matured, offers the potential for fabrication of lightweight, durable, and cost-effective advanced rocket propulsion components.
Brockmeyer, Jerry W., Tuffias, Robert H., Williams, Brian E.
Interfacing a General Purpose Fluid Network Flow Program with the SINDA/G Thermal Analysis Program1999-01-21627/12/1999
A general purpose, one dimensional fluid flow code is currently being interfaced with the thermal analysis program SINDA/G. The flow code, the Generalized Fluid System Simulation Program (GFSSP), is capable of analyzing steady state and transient flow in a complex network. The flow code is capable of modeling several physical phenomena including compressibility effects, phase changes, body forces (such as gravity and centrifugal) and mixture thermodynamics for multiple species. The addition of GFSSP to SINDA/G provides a significant improvement in convective heat transfer modeling for SINDA/G. The interface development is conducted in multiple phases. This paper describes the first phase of the interface which allows for steady and quasi-steady (unsteady solid, steady fluid) conjugate heat transfer modeling.
Schallhorn, Paul, Popok, Dan
Integrated Capillary Engine - Conceptual Overview1999-01-20927/12/1999
A scoping thermal analysis was done to evaluate the general feasibility of capillary pumped heat engines. The analysis was motivated by recent advances in nanoscale materials science that have made it increasingly practical to manufacture high porosity wicks with a median pore diameter on the order of a few nanometers. Capillary pumped heat engines are shown to be generally feasible for wick evaporation rates equivalent to about 1 watt per square centimeter when wick material thermal conductivity on the order of a few W/m-K is assumed. A compact heat engine architecture, referred to as an integrated capillary engine, is introduced.
Williams, Rube B., Kobayashi, Yasunori
Does it Exist, the New Theoretical Limits for Specific Impulse of Rocket Engines94212210/1/1994
The technical thermodynamics solution of rocket engine cycle optimization is introduced here. The traditional approximations of thrust production process are critically evaluated and, as a consequence, a noncontroversial thermodynamic approach is presented. This paper underlines the uniqueness of the physical nature of such a thrust production process. The thermodynamics analysis results show that new limits of theoretical rocket engine specific impulse exist within this approach, and this can lead in future to a reduction of launch costs, environmental benefits for space system use and development of a specific, more efficient rocket system, such as nuclear and SSTO. New investigations ought to be organised and directed toward finding the practical possibilities to implement the ideas introduced here.
Zakharov, Alexander M.
Radioisotope Thermoelectric Generator Waste Heat System for the Cassini Propulsion Module9412696/1/1994
This paper describes the Radioisotope Thermoelectric Generator (RTG) waste heat system and its development, performance, and effectiveness at controlling the temperatures for the propulsion module for the Cassini spacecraft. To explore the Saturn system, the Cassini spacecraft relies on the electrical power from these RTGs, but the large power demand for science and engineering functions severely limits the electrical power available for temperature control. The waste heat system transfers heat radiatively and conductively from each RTG to the propulsion module to maintain temperatures and establish a stable temperature sink to which the rest of the orbiter is thermally anchored.
Mireles, Virgil, Stultz, James W.
The Variable Radioisotope Heater Unit for the Cassini Spacecraft9412686/1/1994
This paper describes the Variable Radioisotope Heater Unit; its development, performance, and effectiveness at controlling the temperatures for the Cassini spacecraft thruster clusters. To explore the Saturn system, the Cassini spacecraft relies on the electrical power from three Radioisotope Thermoelectric Generators, but the large power demand for science and engineering functions severely limits the electrical power available for temperature control purposes. The Variable Radioisotope Heater Unit combines the heating- and temperature-control functions into one nonelectrical self-controlled unit, thereby freeing up electrical heater power for other uses on the spacecraft.
Lyra, Jacqueline C., Stultz, James W.
Carbon Phenolic Ablation Testing and Analysis for High Speed Earth Return9413696/1/1994
Testing of carbon phenolic heatshield materials suitable for the forward heatshield of a high speed Earth return capsule is addressed. Critical aspects of the testing included thermocouple size, orientation and placement, and heat flux calibration to provide accurate data. Post-test examination provides evidence of char and pyrolysis depth, delamination and cracking. Analysis indicates the specific heat capacity has a heating rate dependency, this has been accounted for in the correlation of conductivity with temperature. A fissure model to account for delamination of the samples needed to be included.
Mariage, E., Lacoste, M.
Rocket Based Combined Cycles for Single Stage Rocket9411664/1/1994
Two concepts of rocket based combined cycle for single stage rocket are examined. Concept of SSR power plant consisted of liquid rocket engine and liquid air cycle engine is proposed and substantiated. The advantage of LACE with mixed air+oxygen oxidizer over pure LACE is shown. This combination provides increase of payload fraction of SSR up to 2.7% of the vehicle initial mass compared to rocket with pure LRE. Performance and specific mass of the combined engine are determined. Optimal range of oxidizer/hydrogen mixture ratio is shown. Methods of liquefaction ratio increase are discussed. Second concept considers LRE cycle integration with atmospheric air separation system. The main feature of the scheme is that air separation system provides positive thrust at the initial acceleration. With the same assumption for mass estimation, the concept shows close results to LACE/LRE combination in terms of rocket payload fraction.
Balepin, Vladimir V., Yoshida, Makoto, Kamijo, Kenjiro
Electrohydraulic ServovalvesARP490 (Historical)9/16/1993
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 method 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; and c. provide guidelines for its application on programs which have specified reliability requirements.ility requirements.
A-6B1 Hydraulic Servo Actuation Committee
Optical Methods for the Measurement of the Temperature and Concentration Distribution in Combustion Chambers9324689/1/1993
Two kinds of optical methods for the measurement of the temperature and concentration distribution in combustion chambers by means of the real-time laser holographic interferometry are presented in this paper, which are respectively the one-wavelength laser holographic interferometric technique and the two-wavelength laser holographic interferometric technique. The basic measurement principles of these laser holographic interferometric techniques, the characteristic features of their optical arrangements, the experimental device and testing procedure, and the formula of evaluation of holographic interferograms are described in detail. The experimental results show that the one-wavelength laser holographic interferometric technique is suited to research the temperature distribution and its variation processes in the combustion chambers in detail after combined with a high speed camera, and the two-wavelength laser holographic interferometric technique is suited to investigate the coupled temperature and concentration distribution in the combustion chambers during the period of the fuel evaporation. MUCH RESEARCH WORK has focusd on the investigation of the temperature and concentration distribution in the fuel evaporation process in order to ascertain the formation process of the mixture of air fuel-vapor in the combustion chambers in diesel engines and to obtain the good performance of diesel engines [1]*. Many methods have been used for many years in measuring the temperature and concentration distribution [2]. For example, the most common method is that the temperature distribution in a combustion chamber of a diesel engine is measured by using the thermocouples. However the measurement results from thermocouples are unsatisfactoried because of the poor response property, the effects on the flow field in the combustion chamber, and the point by point measurement information. Some optical methods, such as Schlieren and Shadowgraph methods, can not be used for the quantitative measurements although they can be used for the observation of the in-cylinder process and can obtain the information about a whole field by the evaluation of photographs. In Mach-Zehnder interferometry, the imperfections of windows, lenses and mirrors will influence the accuracy of measurement, and the adjustment of equipment is difficult. Since the laser is invented, the laser holographic interferometic techniques have quickly been developed and have found wide applications. In contrast to conventional interferometric methods, the laser holographic interferometric techniqes offer the considerable advantages of greater experimental and technical simplisity, which means that the necessary optical arrangements are much cheaper; the time needed for adjusting equipment and taking measurements is also greatly reduced; high quality windows, lenses and mirrors are not necessary. In addition, the measurements can be made even in high pressure and high temperature as well as very fast variation process. In this paper two kinds of optical methods for measurement of temperature and concentration distribution in the combustion chambers are presented, which are respectively the one-wavelength laser holographic interferometric technique and the two-wavelength laser holographic interferometric technique. The basic measurement principles of these laser holographic interferometric techniques, the characteristic features of their optical arrangements, the experimental device and testing procedure, and the formula of evaluation of the holographic interferograms are described in detail. The research work done in this paper provides, the author hope, a solid foundation for the investigation of the temperature and concentration distribution in the fuel evaporation process and for further investigation of the formation process of the air fuel-vapor in the combustion chambers in diesel engines.
Zhengbai, Liu
Liquid-Propellant Rocket Engine Testing at Arnold Engineering Development Center9314474/1/1993
A continuing need exists for facilities to test both storable and cryogenic liquid-propellant rocket engines and stages at simulated altitude as part of a responsible acquisition risk reduction program. Storable propellant Intercontinental Ballistic Missile (ICBM) Post Boost Vehicles (PBV) require simulated altitude testing as part of the Aging and Surveillance programs designed to ensure an effective and reliable missile system. Likewise, simulated altitude testing is necessary acquisition risk reduction for advanced cryogenic rocket engines and stages that are being upgraded or developed to satisfy a variety of defense and commercial payload requirements. A review of the liquid rocket test facilities at Arnold Engineering Development Center (AEDC) is presented. The facility capabilities used in support of acquisition risk reduction are described, as are new facility capabilities recently completed or funded. Furthermore, a description of the technology applications available at AEDC in support of liquid rocket diagnostics, analysis, and evaluation techniques is presented.
DeHoff, Bryan, Tucker, Edgar K.
Rocket Engine Test Facilities at Stennis Space Center9314344/1/1993
Ground testing of rocket engines and rocket engine components has played an essential role in the development of flight-rated propulsion systems. The government has built many unique facilities to perform this function. Outlined in this paper are the ground testing capabilities of the National Aeronautics and Space Administration (NASA) John C. Stennis Space Center (SSC), ranging from the component level to fully-integrated, multi-engine stage testing.
Gilbrech, Richard J.
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