Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “liquid rocket”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 757 records · Page 42
All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

Small Signal Modeling and Control of the Hydrogen Mixer for Facility E1

We have undertaken the theoretical modelling of an existing liquid hydrogen (LH2) and gas hydrogen (GH2) mixer subsystem of the E1 Ground Test Facility at NASA John C. Stennis Space Center. The E1 test facility carries out comprehensive ground-based testing and certification of various liquid rocket engines and their components. The mixer described in this work is responsible for combining high pressure before it is fed into a test article. The desired properties are maintained by precise control of the mixture of LH2 and GH2 flows. The mixer is modelled as a general multi-flow lumped volume for single constituent fluids using density and internal energy as states. The set of nonlinear differential equations is linearized about an equilibrium point and the resulting two-state, 3-input linear model is analyzed as a possible candidate for control design.

Barbieri, Enrique↗
NASA NTRS2005

RP-1 Thermal Stability and Copper Based Materials Compatibility Study

A series of electrically heated tube tests was performed at the NASA Glenn Research Center s Heated Tube Facility to investigate the effect that sulfur content, test duration, and tube material play in the overall thermal stability and materials compatibility characteristics of RP-1. Scanning-electron microscopic (SEM) analysis in conjunction with energy dispersive spectroscopy (EDS) were used to characterize the condition of the tube inner wall surface and any carbon deposition or corrosion formed during these runs. Results of the parametric study indicate that tests with standard RP-1 (total sulfur -23 ppm) and pure copper tubing are characterized by a depostion/deposit shedding process producing local wall temperature swings as high as 500 F. The effect of this shedding is to keep total carbon deposition levels relatively constant for run times from 20 minutes up to 5 hours, though increasing tube pressure drops were observed in all runs. Reduction in the total sulfur content of the fuel from 23 ppm to less than 0.1 ppm resulted in the elimination of deposit shedding, local wall temperature variation, and the tube pressure drop increases that were observed in standard sulfur level RP-1 tests. The copper alloy GRCop-84, a copper alloy developed specifically for high heat flux applications, was found to exhibit higher carbon deposition levels compared to identical tests performed in pure copper tubes. Results of the study are consistent with previously published heated tube data which indicates that small changes in fuel total sulfur content can lead to significant differences in the thermal stability of kerosene type fuels and their compatibility with copper based materials. In conjunction with the existing thermal stability database, these findings give insight into the feasibility of cooling a long life, high performance, high-pressure liquid rocket combustor and nozzle with RP-1.

Stiegemeier, B. R.↗
NASA NTRS2005

Matlab GUI for a Fluid Mixer

The Test and Engineering Directorate at NASA John C. Stennis Space Center developed an interest to study the modeling, evaluation, and control of a liquid hydrogen (LH2) and gas hydrogen (GH2) mixer subsystem of a ground test facility. This facility carries out comprehensive ground-based testing and certification of liquid rocket engines including the Space Shuttle Main engine. A software simulation environment developed in MATLAB/SIMULINK (M/S) will allow NASA engineers to test rocket engine systems at relatively no cost. In the progress report submitted in February 2004, we described the development of two foundation programs, a reverse look-up application using various interpolation algorithms, a variety of search and return methods, and self-checking methods to reduce the error in returned search results to increase the functionality of the program. The results showed that these efforts were successful. To transfer this technology to engineers who are not familiar with the M/S environment, a four-module GUI was implemented allowing the user to evaluate the mixer model under open-loop and closed-loop conditions. The progress report was based on an udergraduate Honors Thesis by Ms. Jamie Granger Austin in the Department of Electrical Engineering and Computer Science at Tulane University, during January-May 2003, and her continued efforts during August-December 2003. In collaboration with Dr. Hanz Richter and Dr. Fernando Figueroa we published these results in a NASA Tech Brief due to appear this year. Although the original proposal in 2003 did not address other components of the test facility, we decided in the last few months to extend our research and consider a related pressurization tank component as well. This report summarizes the results obtained towards a Graphical User Interface (GUI) for the evaluation and control of the hydrogen mixer subsystem model and for the pressurization tank each taken individually. Further research would combine the two components - mixer and tank, for a more realistic simulation tool.

Barbieri, Enrique↗
NASA NTRS2004

RTF Analysis-Propulsion/Flowliner

The objective of the current task is to provide a computational framework for design and analysis of the entire fuel supply system of a liquid rocket engine, including high fidelity unsteady turbopump flow analysis. Time-accurate results obtained from shuttle fuel flowliner analysis using 66 Million grid points with 262 overlapped zones will be shown. We will present analysis results along with performance data of the simulation runs on Supercomputers such as Columbia.

Kiris, Centin↗
NASA NTRS2004

Investigation of Navier-Stokes Code Verification and Design Optimization

With rapid progress made in employing computational techniques for various complex Navier-Stokes fluid flow problems, design optimization problems traditionally based on empirical formulations and experiments are now being addressed with the aid of computational fluid dynamics (CFD). To be able to carry out an effective CFD-based optimization study, it is essential that the uncertainty and appropriate confidence limits of the CFD solutions be quantified over the chosen design space. The present dissertation investigates the issues related to code verification, surrogate model-based optimization and sensitivity evaluation. For Navier-Stokes (NS) CFD code verification a least square extrapolation (LSE) method is assessed. This method projects numerically computed NS solutions from multiple, coarser base grids onto a freer grid and improves solution accuracy by minimizing the residual of the discretized NS equations over the projected grid. In this dissertation, the finite volume (FV) formulation is focused on. The interplay between the xi concepts and the outcome of LSE, and the effects of solution gradients and singularities, nonlinear physics, and coupling of flow variables on the effectiveness of LSE are investigated. A CFD-based design optimization of a single element liquid rocket injector is conducted with surrogate models developed using response surface methodology (RSM) based on CFD solutions. The computational model consists of the NS equations, finite rate chemistry, and the k-6 turbulence closure. With the aid of these surrogate models, sensitivity and trade-off analyses are carried out for the injector design whose geometry (hydrogen flow angle, hydrogen and oxygen flow areas and oxygen post tip thickness) is optimized to attain desirable goals in performance (combustion length) and life/survivability (the maximum temperatures on the oxidizer post tip and injector face and a combustion chamber wall temperature). A preliminary multi-objective optimization study is carried out using a geometric mean approach. Following this, sensitivity analyses with the aid of variance-based non-parametric approach and partial correlation coefficients are conducted using data available from surrogate models of the objectives and the multi-objective optima to identify the contribution of the design variables to the objective variability and to analyze the variability of the design variables and the objectives. In summary the present dissertation offers insight into an improved coarse to fine grid extrapolation technique for Navier-Stokes computations and also suggests tools for a designer to conduct design optimization study and related sensitivity analyses for a given design problem.

Vaidyanathan, Rajkumar↗
NASA NTRS2005

Surrogate-based Analysis and Optimization

A major challenge to the successful full-scale development of modem aerospace systems is to address competing objectives such as improved performance, reduced costs, and enhanced safety. Accurate, high-fidelity models are typically time consuming and computationally expensive. Furthermore, informed decisions should be made with an understanding of the impact (global sensitivity) of the design variables on the different objectives. In this context, the so-called surrogate-based approach for analysis and optimization can play a very valuable role. The surrogates are constructed using data drawn from high-fidelity models, and provide fast approximations of the objectives and constraints at new design points, thereby making sensitivity and optimization studies feasible. This paper provides a comprehensive discussion of the fundamental issues that arise in surrogate-based analysis and optimization (SBAO), highlighting concepts, methods, techniques, as well as practical implications. The issues addressed include the selection of the loss function and regularization criteria for constructing the surrogates, design of experiments, surrogate selection and construction, sensitivity analysis, convergence, and optimization. The multi-objective optimal design of a liquid rocket injector is presented to highlight the state of the art and to help guide future efforts.

Queipo, Nestor V.↗
NASA NTRS2005

Gas-Centered Swirl Coaxial Liquid Injector Evaluations

Development of Liquid Rocket Engines is expensive. Extensive testing at large scales usually required. In order to verify engine lifetime, large number of tests required. Limited Resources available for development. Sub-scale cold-flow and hot-fire testing is extremely cost effective. Could be a necessary (but not sufficient) condition for long engine lifetime. Reduces overall costs and risk of large scale testing. Goal: Determine knowledge that can be gained from sub-scale cold-flow and hot-fire evaluations of LRE injectors. Determine relationships between cold-flow and hot-fire data.

Cohn, A. K.↗
NASA NTRS2006

Experimental Methodology for Measuring Combustion and Injection-Coupled Responses

A Russian scaling methodology for liquid rocket engines utilizing a single, full scale element is reviewed. The scaling methodology exploits the supercritical phase of the full scale propellants to simplify scaling requirements. Many assumptions are utilized in the derivation of the scaling criteria. A test apparatus design is presented to implement the Russian methodology and consequently verify the assumptions. This test apparatus will allow researchers to assess the usefulness of the scaling procedures and possibly enhance the methodology. A matrix of the apparatus capabilities for a RD-170 injector is also presented. Several methods to enhance the methodology have been generated through the design process.

Cavitt, Ryan C.↗
NASA NTRS2009

Liquid Engine Design: Effect of Chamber Dimensions on Specific Impulse

Which assumption of combustion chemistry - frozen or equilibrium - should be used in the prediction of liquid rocket engine performance calculations? Can a correlation be developed for this? A literature search using the LaSSe tool, an online repository of old rocket data and reports, was completed. Test results of NTO/Aerozine-50 and Lox/LH2 subscale and full-scale injector and combustion chamber test results were found and studied for this task. NASA code, Chemical Equilibrium with Applications (CEA) was used to predict engine performance using both chemistry assumptions, defined here. Frozen- composition remains frozen during expansion through the nozzle. Equilibrium- instantaneous chemical equilibrium during nozzle expansion. Chamber parameters were varied to understand what dimensions drive chamber C* and Isp. Contraction Ratio is the ratio of the nozzle throat area to the area of the chamber. L is the length of the chamber. Characteristic chamber length, L*, is the length that the chamber would be if it were a straight tube and had no converging nozzle. Goal: Develop a qualitative and quantitative correlation for performance parameters - Specific Impulse (Isp) and Characteristic Velocity (C*) - as a function of one or more chamber dimensions - Contraction Ratio (CR), Chamber Length (L ) and/or Characteristic Chamber Length (L*). Determine if chamber dimensions can be correlated to frozen or equilibrium chemistry.

Hoggard, Lindsay↗
NASA NTRS2009

Vibration Challenges in the Design of NASA's Ares Launch Vehicles

This paper focuses on the vibration challenges inherent in the design of NASA s Ares launch vehicles. A brief overview of the launch system architecture is provided to establish the context for the discussion. Following this is a general discussion of the design considerations and analytical disciplines that are affected by vibration. The first challenge discussed is that of coupling between the vehicle flight control system and fundamental vibrational modes of the vehicle. The potential destabilizing influence of the vibrational dynamics is described along with discussion of the typical methods employed to overcome this issue. Next is a general discussion of the process for developing the design loads for the primary structure. This includes quasi-steady loads and dynamic loads induced by the structural dynamic response. The two principal parts of this response are the gust induced responses of the lower frequency modes and the buffet induced responses of the higher frequency modes. Structural dynamic model validation will also be addressed. Following this, discussions of three somewhat unique topics of Pogo Instability, Solid Booster Thrust Oscillation, and Liquid Rocket Engine Turbopump Rotordynamic Stability and Response are presented.

Ryan, Stephen G.↗
NASA NTRS2010

Planning for Plume Diagnostics for Ground Testing of J-2X Engines at the SSC

John C. Stennis Space Center (SSC) is the premier test facility for liquid rocket engine development and certification for the National Aeronautics and Space Administration (NASA). Therefore, it is no surprise that the SSC will play the most prominent role in the engine development testing and certification for the J-2X engine. The Pratt & Whitney Rocketdyne J-2X engine has been selected by the Constellation Program to power the Ares I Upper Stage Element and the Ares V Earth Departure Stage in NASA s strategy of risk mitigation for hardware development by building on the Apollo program and other lessons learned to deliver a human-rated engine that is on an aggressive development schedule, with first demonstration flight in 2010 and human test flights in 2012. Accordingly, J-2X engine design, development, test, and evaluation is to build upon heritage hardware and apply valuable experience gained from past development and testing efforts. In order to leverage SSC s successful and innovative expertise in the plume diagnostics for the space shuttle main engine (SSME) health monitoring,1-10 this paper will present a blueprint for plume diagnostics for various proposed ground testing activities for J-2X at SSC. Complete description of the SSC s test facilities, supporting infrastructure, and test facilities is available in Ref. 11. The A-1 Test Stand is currently being prepared for testing the J-2X engine at sea level conditions. The A-2 Test Stand is currently being used for testing the SSME and may also be used for testing the J-2X engine at sea level conditions in the future. Very recently, ground-breaking ceremony for the new A-3 rocket engine test stand took place at SSC on August 23, 2007. A-3 is the first large - scale test stand to be built at the SSC since the A and B stands were constructed in the 1960s. The A-3 Test Stand will be used for testing J-2X engines under vacuum conditions simulating high altitude operation at approximately 30,480 m (100,000 ft). To achieve the simulated altitude environment, chemical steam generators using isopropyl alcohol, LOX, and RELEASED - Printed documents may be obsolete; validate prior to use. water would run for the duration of the test and would generate approximately 2096 Kg/s of steam to reduce pressure in the test cell and downstream of the engine. The testing at the A-3 Test Stand is projected to begin in late 2010, meanwhile the J-2X component testing on A-1 is scheduled to begin later this year.

SaintCyr, William W.↗
NASA NTRS2010

A Framework for Integrated Component and System Analyses of Instabilities

Instabilities associated with fluid handling and operation in liquid rocket propulsion systems and test facilities usually manifest themselves as structural vibrations or some form of structural damage. While the source of the instability is directly related to the performance of a component such as a turbopump, valve or a flow control element, the associated pressure fluctuations as they propagate through the system have the potential to amplify and resonate with natural modes of the structural elements and components of the system. In this paper, the authors have developed an innovative multi-level approach that involves analysis at the component and systems level. The primary source of the unsteadiness is modeled with a high-fidelity hybrid RANS/LES based CFD methodology that has been previously used to study instabilities in feed systems. This high fidelity approach is used to quantify the instability and understand the physics associated with the instability. System response to the driving instability is determined through a transfer matrix approach wherein the incoming and outgoing pressure and velocity fluctuations are related through a transfer (or transmission) matrix. The coefficients of the transfer matrix for each component (i.e. valve, pipe, orifice etc.) are individually derived from the flow physics associated with the component. A demonstration case representing a test loop/test facility comprised of a network of elements is constructed with the transfer matrix approach and the amplification of modes analyzed as the instability propagates through the test loop.

Ahuja, Vineet↗
NASA NTRS2010

Nonlinear Combustion Instability Prediction

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.

Flandro, Gary↗
NASA NTRS2010

Liquid Propulsion: Propellant Feed System Design

The propellant feed system of a liquid rocket engine determines how the propellants are delivered from the tanks to the thrust chamber. They are generally classified as either pressure fed or pump fed. The pressure-fed system is simple and relies on tank pressures to feed the propellants into the thrust chamber. This type of system is typically used for space propulsion applications and auxiliary propulsion applications requiring low system pressures and small quantities of propellants. In contrast, the pump-fed system is used for high pressure, high performance applications. The selection of one propellant feed system over another is determined based on design trade studies at both the engine and vehicle levels. This chapter first provides a brief overview of the basic configurations of pressure-fed systems. Pump-fed systems are then discussed with greater detail given to the turbomachinery design. Selected design requirements and configurations are provided.

Cannon, James L.↗
NASA NTRS2010

RP-2 Thermal Stability and Heat Transfer Investigation for Hydrocarbon Boost Engines

A series of electrically heated tube tests were performed at the NASA Glenn Research Center s Heated Tube Facility to investigate the use of RP-2 as a fuel for next generation regeneratively cooled hydrocarbon boost engines. The effect that test duration, operating condition and test piece material have on the overall thermal stability and materials compatibility characteristics of RP-2 were evaluated using copper and 304 stainless steel test sections. The copper tests were run at 1000 psia, heat flux up to 6.0 Btu/in.2-sec, and wall temperatures up to 1180 F. Preliminary results, using measured wall temperature as an indirect indicator of the carbon deposition process, show that in copper test pieces above approximately 850 F, RP-2 begins to undergo thermal decomposition resulting in local carbon deposits. Wall temperature traces show significant local temperature increases followed by near instantaneous drops which have been attributed to the carbon deposition/shedding process in previous investigations. Data reduction is currently underway for the stainless steel test sections and carbon deposition measurements will be performed in the future for all test sections used in this investigation. In conjunction with the existing thermal stability database, these findings give insight into the feasibility of cooling a long life, high performance, high-pressure liquid rocket combustor and nozzle with RP-2.

VanNoord, J. L.↗
NASA NTRS2011

Pros, Cons, and Alternatives to Weight Based Cost Estimating

Many cost estimating tools use weight as a major parameter in projecting the cost. This is often combined with modifying factors such as complexity, technical maturity of design, environment of operation, etc. to increase the fidelity of the estimate. For a set of conceptual designs, all meeting the same requirements, increased weight can be a major driver in increased cost. However, once a design is fixed, increased weight generally decreases cost, while decreased weight generally increases cost - and the relationship is not linear. Alternative approaches to estimating cost without using weight (except perhaps for materials costs) have been attempted to try to produce a tool usable throughout the design process - from concept studies through development. This paper will address the pros and cons of using weight based models for cost estimating, using liquid rocket engines as the example. It will then examine approaches that minimize the imp~ct of weight based cost estimating. The Rocket Engine- Cost Model (RECM) is an attribute based model developed internally by Pratt & Whitney Rocketdyne for NASA. RECM will be presented primarily to show a successful method to use design and programmatic parameters instead of weight to estimate both design and development costs and production costs. An operations model developed by KSC, the Launch and Landing Effects Ground Operations model (LLEGO), will also be discussed.

Joyner, Claude R.↗
NASA NTRS2004

Calculating Mass Diffusion in High-Pressure Binary Fluids

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.

Bellan, Josette↗
NASA NTRS2003

Update on Area Production in Mixing of Supercritical Fluids

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 gas-turbine 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.

Okongo, Nora↗