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NASA NTRS2003

Numerical Modeling of Pulse Detonation Rocket Engine Gasdynamics and Performance

This paper presents viewgraphs on the numerical modeling of pulse detonation rocket engines (PDRE), with an emphasis on the Gasdynamics and performance analysis of these engines. The topics include: 1) Performance Analysis of PDREs; 2) Simplified PDRE Cycle; 3) Comparison of PDRE and Steady-State Rocket Engines (SSRE) Performance; 4) Numerical Modeling of Quasi 1-D Rocket Flows; 5) Specific PDRE Geometries Studied; 6) Time-Accurate Thrust Calculations; 7) PDRE Performance (Geometries A B C and D); 8) PDRE Blowdown Gasdynamics (Geom. A B C and D); 9) PDRE Geometry Performance Comparison; 10) PDRE Blowdown Time (Geom. A B C and D); 11) Specific SSRE Geometry Studied; 12) Effect of F-R Chemistry on SSRE Performance; 13) PDRE/SSRE Performance Comparison; 14) PDRE Performance Study; 15) Grid Resolution Study; and 16) Effect of F-R Chemistry on SSRE Exit Species Mole Fractions.

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NASA NTRS2003

Transient Simulation of the Integrated Powerhead Demonstrator (IPD) Rocket Engine

The IPD rocket engine is the product of a joint Air Force/NASA program to demonstrate the concept of a full-flow staged combustion power cycle. In this type of rocket engine cycle, both the entire fuel flow and the entire oxidizer flow are combusted in one of two pre-bumers and used to drive two turbopumps, thus utilizing the flow work of the total propellant flow. The basis of the IPD engine program relies on integrating newly developed and pre-existing hardware to demonstrate the component and material technologies to make this concept feasible, while simultaneously saving development time and costs. To provide insight to the project team and contractors during engine development and test phases, the engine system was modeled at MSFC using the Rocket Engine Transient Simulation (ROCETS) software to analyze system performance and determine component integration issues. The ROCETS software is used extensively a MSFC to perform steady-state power-balances and transient simulations of thermodynamic power and general fluid systems. The software is favored for its capability to solve large systems of non-linear equations, its librarie of fluid properties and flow devices, its flexibility to modify existing code to improve the physics-derived approximations of real fluid thermodynamic behavior, and the ability to add unique system constraints. The purpose of this paper is to present the methodology used to model the IPD engine system, detail the pitfalls encountered with the software, and explain the approximations made to more accurately represent engine component and fluid behavior. Engine system performance output from the model will be presented and explained in comparison with real fluid and component behavior.

Guidos, Mike↗

NASA NTRS1951

Combustion Instability in an Acid-Heptane Rocket with a Pressurized-Gas Propellant Pumping System

Results of experimental measurements of low-frequency combustion instability of a 300-pound thrust acid-heptane rocket engine were compared to the trends predicted by an analysis of combustion instability in a rocket engine with a pressurized-gas propellant pumping system. The simplified analysis, which assumes a monopropellant model, was based on the concept of a combustion the delay occurring from the moment of propellant injection to the moment of propellant combustion. This combustion time delay was experimentally measured; the experimental values were of approximately half the magnitude predicted by the analysis. The pressure-fluctuation frequency for a rocket engine with a characteristic length of 100 inches and operated at a combustion-chamber pressure of 280 pounds per square inch absolute was 38 cycles per second; the analysis indicated. a frequency of 37 cycles per second. Increasing combustion-chamber characteristic length decreased the pressure-fluctuation frequency, in conformity to the analysis. Increasing the chamber operating pressure or increasing the injector pressure drop increased the frequency. These latter two effects are contrary to the analysis; the discrepancies are attributed to the conflict between the assumptions made to simplify the analysis and the experimental conditions. Oxidant-fuel ratio had no apparent effect on the experimentally measured pressure-fluctuation frequency for acid-heptane ratios from 3.0 to 7.0. The frequencies decreased with increased amplitude of the combustion-chamber pressure variations. The analysis indicated that if the combustion time delay were sufficiently short, low-frequency combustion instability would be eliminated.

Tischler, Adelbert O.↗

NASA NTRS1947

Report on Rocket Power Plants Based on T-Substance

In the search for an energy source independent of air for the propulsion of underwater craft, attention was early concentrated on T-substance. It was possible to convince the OKM [NACA comment: Navy High Command] very quickly of the importance of this material. In 1934, the first experiments were undertaken. A difficulty was at once presented by the limited concentration that had been attained. At first only 60 percent T-substance could be supplied; this amount was later increased to as much as 85 percent. Decomposition and combustion experiments conducted on the grounds of the CPVA in Kiel-Dietrichsdorf led to the first practical information as to the technical feasibility of the use of T-substance. New perspectives soon developed because a method of concentrated energy production had been found here, which was capable of many applications. The idea of using this energy for the propulsion of missiles either in guns or as rockets suggested itself and appropriate proposals, which quickly led to the construction of the first experimental devices, were made to the official quarters concerned. In January 1937, the first flight of a DVL aircraft with T-substance auxiliary propulsion took place at Alimbsmuhle in the presence of Colonel Udet, who piloted the third flight. In June 1937, the first T-substance rockets were fired (Altenwalde). Then in rapid succession take-off auxiliary, main propulsion, and other rocket drives were brought out in experimental versions. Hydrogen peroxide is a well known chemical, which is widely used in the textile industry. Its chemical and physical properties as well as the processes of manufacture and use are familiar and have been described in voluminous books and papers, Nevertheless, much developmental work was required to open the way for T-substance as a usable oxygen carrier. In fact the utilization of hydrogen peroxide as the oxygen carrier for energy production had hitherto been the subject only of isolated suggestions, which have never been developed beyond the stage of theoretical discussion.

Walter, Hellmuth↗

NASA NTRS1955

Low-Lift Drag of the Grumman F9F-9 Airplane as Obtained by a 1/7.5-Scale Rocket-Boosted Model and by Three 1/45.85-Scale Equivalent-Body Models between Mach Numbers of 0.8 and 1.3, TED No. NACA DE 391

Low-lift drag data are presented herein for one 1/7.5-scale rocket-boosted model and three 1/45.85-scale equivalent-body models of the Grumman F9F-9 airplane, The data were obtained over a Reynolds number range of about 5 x 10(exp 6) to 10 x 10(exp 6) based on wing mean aerodynamic chord for the rocket model and total body length for the equivalent-body models. The rocket-boosted model showed a drag rise of about 0,037 (based on included wing area) between the subsonic level and the peak supersonic drag coefficient at the maximum Mach number of this test. The base drag coefficient measured on this model varied from a value of -0,0015 in the subsonic range to a maximum of about 0.0020 at a Mach number of 1.28, Drag coefficients for the equivalent-body models varied from about 0.125 (based on body maximum area) in the subsonic range to about 0.300 at a Mach number of 1.25. Increasing the total fineness ratio by a small amount raised the drag-rise Mach number slightly.

Stevens, Joseph E.↗

NASA NTRS2005

Modal Survey of ETM-3, A 5-Segment Derivative of the Space Shuttle Solid Rocket Booster

The complex interactions between internal motor generated pressure oscillations and motor structural vibration modes associated with the static test configuration of a Reusable Solid Rocket Motor have potential to generate significant dynamic thrust loads in the 5-segment configuration (Engineering Test Motor 3). Finite element model load predictions for worst-case conditions were generated based on extrapolation of a previously correlated 4-segment motor model. A modal survey was performed on the largest rocket motor to date, Engineering Test Motor #3 (ETM-3), to provide data for finite element model correlation and validation of model generated design loads. The modal survey preparation included pretest analyses to determine an efficient analysis set selection using the Effective Independence Method and test simulations to assure critical test stand component loads did not exceed design limits. Historical Reusable Solid Rocket Motor modal testing, ETM-3 test analysis model development and pre-test loads analyses, as well as test execution, and a comparison of results to pre-test predictions are discussed.

Nielsen, D.↗

NASA NTRS2005

Liquid Rocket Engine Testing Overview

Contents include the following: Objectives and motivation for testing. Technology, Research and Development Test and Evaluation (RDT&E), evolutionary. Representative Liquid Rocket Engine (LRE) test compaigns. Apollo, shuttle, Expandable Launch Vehicles (ELV) propulsion. Overview of test facilities for liquid rocket engines. Boost, upper stage (sea-level and altitude). Statistics (historical) of Liquid Rocket Engine Testing. LOX/LH, LOX/RP, other development. Test project enablers: engineering tools, operations, processes, infrastructure.

Rahman, Shamim↗

NASA NTRS2000

Thermal Barriers Developed for Solid Rocket Motor Nozzle Joints

Space shuttle solid rocket motor case assembly joints are sealed with conventional O-ring seals that are shielded from 5500 F combustion gases by thick layers of insulation and by special joint-fill compounds that fill assembly splitlines in the insulation. On a number of occasions, NASA has observed hot gas penetration through defects in the joint-fill compound of several of the rocket nozzle assembly joints. In the current nozzle-to-case joint, NASA has observed penetration of hot combustion gases through the joint-fill compound to the inboard wiper O-ring in one out of seven motors. Although this condition does not threaten motor safety, evidence of hot gas penetration to the wiper O-ring results in extensive reviews before resuming flight. The solid rocket motor manufacturer (Thiokol) approached the NASA Glenn Research Center at Lewis Field about the possibility of applying Glenn's braided fiber preform seal as a thermal barrier to protect the O-ring seals. Glenn and Thiokol are working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design and by using a braided carbon fiber thermal barrier that would resist any hot gases that the J-leg does not block.

Steinetz, Bruce M.↗

NASA NTRS2005

Rocket Engine Nozzle Side Load Transient Analysis Methodology: A Practical Approach

At the sea level, a phenomenon common with all rocket engines, especially for a highly over-expanded nozzle, during ignition and shutdown is that of flow separation as the plume fills and empties the nozzle, Since the flow will be separated randomly. it will generate side loads, i.e. non-axial forces. Since rocket engines are designed to produce axial thrust to power the vehicles, it is not desirable to be excited by non-axial input forcing functions, In the past, several engine failures were attributed to side loads. During the development stage, in order to design/size the rocket engine components and to reduce the risks, the local dynamic environments as well as dynamic interface loads have to be defined. The methodology developed here is the way to determine the peak loads and shock environments for new engine components. In the past it is not feasible to predict the shock environments, e.g. shock response spectra, from one engine to the other, because it is not scaleable. Therefore, the problem has been resolved and the shock environments can be defined in the early stage of new engine development. Additional information is included in the original extended abstract.

Shi, John J.↗

NASA NTRS2005

Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

Pulse detonation rocket engines (PDREs) have generated research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred considerable interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the progress that has been made in comparing the available experimental measurements with analytical and numerical models. In recent work by the author, a quasi-one-dimensional, finite rate chemistry CFD model was utilized to study the gasdynamics and performance characteristics of PDREs over a range of blowdown pressure ratios from 1-1000. Models of this type are computationally inexpensive, and enable first-order parametric studies of the effect of several nozzle and extension geometries on PDRE performance over a wide range of conditions. However, the quasi-one-dimensional approach is limited in that it cannot properly capture the multidimensional blast wave and flow expansion downstream of the PDRE, nor can it resolve nozzle flow separation if present. Moreover, the previous work was limited to single-pulse calculations. In this paper, an axisymmetric finite rate chemistry model is described and utilized to study these issues in greater detail. Example Mach number contour plots showing the multidimensional blast wave and nozzle exhaust plume are shown. The performance results are compared with the quasi-one-dimensional results from the previous paper. Both Euler and Navier-Stokes solutions are calculated in order to determine the effect of viscous effects in the nozzle flowfield. Additionally, comparisons of the model results to performance data from CalTech, as well as experimental flowfield measurements from Stanford University, are also reported.

Morris, Christopher I.↗

NASA NTRS2005

AXISYMMETRIC, THROTTLEABLE NON-GIMBALLED ROCKET ENGINE

A rocket engine assembly is provided for a vertically launched rocket vehicle. A rocket engine housing of the assembly includes two or more combustion chambers each including an outlet end defining a sonic throat area. A propellant supply for the combustion chambers includes a throttling injector, associated with each of the combustion chambers and located opposite to sonic throat area, which injects the propellant into the associated combustion chamber. A modulator, which may form part of the injector, and which is controlled by a controller, modulates the flow rate of the propellant to the combustion chambers so that the chambers provide a vectorable net thrust. An expansion nozzle or body located downstream of the throat area provides expansion of the combustion gases produced by the combustion chambers so as to increase the net thrust.

Sackheim, Robert L.↗

NASA NTRS2006

Design Considerations for the Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

Nuclear Thermal Rockets or NTR's have been suggested as a propulsion system option for vehicles traveling to the moon or Mars. These engines are capable of providing high thrust at specific impulses at least twice that of today s best chemical engines. The performance constraints on these engines are mainly the result of temperature limitations on the fuel coupled with a limited ability to withstand chemical attack by the hot hydrogen propellant. To operate at maximum efficiency, fuel forms are desired which can withstand the extremely hot, hostile environment characteristic of NTR operation for at least several hours. The simulation of such an environment would require an experimental device which could simultaneously approximate the power, flow, and temperature conditions which a nuclear fuel element (or partial element) would encounter during NTR operation. Such a simulation would allow detailed studies of the fuel behavior and hydrogen flow characteristics under reactor like conditions to be performed. The goal of these simulations would be directed toward expanding the performance envelope of NTR engines over that which was demonstrated during the Rover and NERVA nuclear rocket programs of the 1970's. Currently, such a simulator is nearing completion at the Marshall Space Flight Center, and will shortly be used in the future to evaluate a wide variety of he1 element designs and the materials of which they are constructed. This present work addresses the initial experimental objectives of the Nuclear Thermal Rocket Element Environmental Simulator or NTREES and some of the design considerations which were considered prior to and during its construction.

Emrich, Bill↗

NASA NTRS2006

Hydrocarbon-Fueled Rocket Engine Plume Diagnostics: Analytical Developments and Experimental Results

A viewgraph presentation describing experimental results and analytical developments about plume diagnostics for hydrocarbon-fueled rocket engines is shown. The topics include: 1) SSC Plume Diagnostics Background; 2) Engine Health Monitoring Approach; 3) Rocket Plume Spectroscopy Simulation Code; 4) Spectral Simulation for 10 Atomic Species and for 11 Diatomic Molecular Electronic Bands; 5) "Best" Lines for Plume Diagnostics for Hydrocarbon-Fueled Rocket Engines; 6) Experimental Set Up for the Methane Thruster Test Program and Experimental Results; and 7) Summary and Recommendations.

Tejwani, Gopal D.↗

NASA NTRS2006

SHARPI/PICTURE Sounding Rocket Telescope

The Solar High Angular Resolution Photometric Imager (SHARPI)/Planet Imaging Concept Testbed Using a Rocket Experiment (PICTURE) Sounding Rocket Telescope is described. The topics include: 1) Lightweight precision mirror development; 2) Two sounding rocket concepts sharing a telescope; 3) Optical Telescope Assembly (OTA) overview; 4) PM development program; 5) PM figure testing; 6) Mirror coatings; 7) PM mount and verification; 8) Secondary Mirror (SM); and 9) OTA.

Content, D.↗

NASA NTRS2006

The Polar Summer MLT Plasma Environment as seen by the DROPPS Sounding Rockets

During early July, 1999, the DROPPS (Distribution and Role of Particles in the Polar Summer Mesosphere) campaign launched two rocket payloads whose purpose was to study the polar summer MLT (mesosphere and lower thermosphere), particularly PMSEs (polar mesospheric summer echoes) and PMCs (polar mesospheric clouds). The rockets were launched from the And(\o)ya Rocket Range in Norway the nights of the 5th and 14th of July. Both payloads included a front-mounted PID (Particle Impact Detector) consisting of charge and mass telescopes to measure aerosol and dust mass distributions. Ice particles of nanometer size are believed to be responsible for PMSEs through the process of electron scavenging. Evidence for this process is suggested, for example, by the presence of an electron "biteout" simultaneously measured by several instruments at an altitude of $\sim$82-87 km during the first DROPPS launch. This presentation will characterize similarities and differences between both flights as seen by the charge and mass telescopes, starting at launch until the loss of data on the downleg of each flight. Various stages of the flights will be considered in detail, such as the PMSE layer and the apogee at 117 km, as well as the calibration of the data before launch.

Assis, Michael P.↗

NASA NTRS2007

Flow Separation Side Loads Excitation of Rocket Nozzle FEM

Modern rocket nozzles are designed to operate over a wide range of altitudes, and are also built with large aspect ratios to enable high efficiencies. Nozzles designed to operate over specific regions of a trajectory are being replaced in modern launch vehicles by those that are designed to operate from earth to orbit. This is happening in parallel with modern manufacturing and wall cooling techniques allowing for larger aspect ratio nozzles to be produced. Such nozzles, though operating over a large range of altitudes and ambient pressures, are typically designed for one specific altitude. Above that altitude the nozzle flow is 'underexpanded' and below that altitude, the nozzle flow is 'overexpanded'. In both conditions the nozzle produces less than the maximum possible thrust at that altitude. Usually the nozzle design altitude is well above sea level, leaving the nozzle flow in an overexpanded state for its start up as well as for its ground testing where, if it is a reusable nozzle such as the Space Shuttle Main Engine (SSME), the nozzle will operate for the majority of its life. Overexpansion in a rocket nozzle presents the critical, and sometimes design driving, problem of flow separation induced side loads. To increase their understanding of nozzle side loads, engineers at MSFC began an investigation in 2000 into the phenomenon through a task entitled "Characterization and Accurate Modeling of Rocket Engine Nozzle Side Loads", led by A. Brown. The stated objective of this study was to develop a methodology to accurately predict the character and magnitude of nozzle side loads. The study included further hot-fire testing of the MC-l engine, cold flow testing of subscale nozzles, CFD analyses of both hot-fire and cold flow nozzle testing, and finite element (fe.) analysis of the MC-1 engine and cold flow tested nozzles. A follow on task included an effort to formulate a simplified methodology for modeling a side load during a two nodal diameter fluid/structure interaction for a single moment in time.

Smalley, Kurt B.↗

NASA NTRS2008

Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

To support the eventual development of a nuclear thermal rocket engine, a state-of-the-art experimental test setup has been constructed to evaluate the performance characteristics of candidate fuel element materials and geometries in representative environments. The test device simulates the environmental conditions (minus the radiation) to which nuclear rocket fuel components will be subjected during reactor operation. Test articles mounted in the simulator are inductively heated in such a manner as to accurately reproduce the temperatures and heat fluxes normally expected to occur as a result of nuclear fission while at the same time being exposed to flowing hydrogen. This project is referred to as the Nuclear Thermal Rocket Element Environment Simulator or NTREES. The NTREES device is located at the Marshall Space flight Center in a laboratory which has been modified to accommodate the high powers required to heat the test articles to the required temperatures and to handle the gaseous hydrogen flow required for the tests. Other modifications to the laboratory include the installation of a nitrogen gas supply system and a cooling water supply system. During the design and construction of the facility, every effort was made to comply with all pertinent regulations to provide assurance that the facility could be operated in a safe and efficient manner. The NTREES system can currently supply up to 50 kW of inductive heating to the fuel test articles, although the facility has been sized to eventually allow test article heating levels of up to several megawatts.

Emrich, William J., Jr.↗

NASA NTRS2006

An Introduction to Rockets - or - Never Leave Geeks Unsupervised

An introduction to rockets along with a brief history Newton's third law is presented. The contents include: 1) What is a Rocket?; 2) A Brief History; 3) Newton's Third Law; 4) A Brief History; 5) Mission Requirements; 6) Some Orbital Measurements; 7) Self Eating Watermelon; 8) Orbital Inclinations; 9) 28.5 Equatorial Orbit; 10) 51.6 Orbit (ISS); 11) Polar Orbit; 12) Geostationary Orbit; 13) Liquid Rocket; 13) Liquids vs. Solids; 14) Liquids; 15) Systems Integration; 16) Integration (NFL!); 17) Guidance Systems; 18) Vectored Thrust; 19) Spin Stabilization; 20) Aerodynamic Stability (Fire Arrows); and 21) Center of Gravity & Center of Pressure.

Mellett, Kevin↗