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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.

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At least 217 records · Page 12
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NASA NTRS2000

Active Combustion Control for Aircraft Gas Turbine Engines

Lean-burning combustors are susceptible to combustion instabilities. Additionally, due to non-uniformities in the fuel-air mixing and in the combustion process, there typically exist hot areas in the combustor exit plane. These hot areas limit the operating temperature at the turbine inlet and thus constrain performance and efficiency. Finally, it is necessary to optimize the fuel-air ratio and flame temperature throughout the combustor to minimize the production of pollutants. In recent years, there has been considerable activity addressing Active Combustion Control. NASA Glenn Research Center's Active Combustion Control Technology effort aims to demonstrate active control in a realistic environment relevant to aircraft engines. Analysis and experiments are tied to aircraft gas turbine combustors. Considerable progress has been shown in demonstrating technologies for Combustion Instability Control, Pattern Factor Control, and Emissions Minimizing Control. Future plans are to advance the maturity of active combustion control technology to eventual demonstration in an engine environment.

DeLaat, John C.↗
NASA NTRS2003

Control of Thermo-Acoustics Instabilities: The Multi-Scale Extended Kalman Approach

"Multi-Scale Extended Kalman" (MSEK) is a novel model-based control approach recently found to be effective for suppressing combustion instabilities in gas turbines. A control law formulated in this approach for fuel modulation demonstrated steady suppression of a high-frequency combustion instability (less than 500Hz) in a liquid-fuel combustion test rig under engine-realistic conditions. To make-up for severe transport-delays on control effect, the MSEK controller combines a wavelet -like Multi-Scale analysis and an Extended Kalman Observer to predict the thermo-acoustic states of combustion pressure perturbations. The commanded fuel modulation is composed of a damper action based on the predicted states, and a tones suppression action based on the Multi-Scale estimation of thermal excitations and other transient disturbances. The controller performs automatic adjustments of the gain and phase of these actions to minimize the Time-Scale Averaged Variances of the pressures inside the combustion zone and upstream of the injector. The successful demonstration of Active Combustion Control with this MSEK controller completed an important NASA milestone for the current research in advanced combustion technologies.

Le, Dzu K.↗
NASA NTRS2015

Advancing the State-of-the-Practice for Liquid Rocket Engine Injector Design

Current shortcomings in both the overall injector design process and its underlying combustion stability assessment methodology are rooted in the use of empirically based or low fidelity representations of complex physical phenomena and geometry details that have first order effects on performance, thermal environments and combustion stability. The result is a design and analysis capability that is often inadequate to reliably arrive at a suitable injector design in an efficient manner. Specifically, combustion instability has been particularly difficult to predict and mitigate. Large hydrocarbon-fueled booster engines have been especially problematic in this regard. Where combustion instability has been a problem, costly and time-consuming redesign efforts have often been an unfortunate consequence. This paper presents an overview of a recently completed effort at NASA Marshall Space Flight Center to advance the state-of-the-practice for liquid rocket engine injector design. Multiple perturbations of a gas-centered swirl coaxial (GCSC) element that burned gaseous oxygen and RP-1 were designed, assessed for combustion stability, and tested. Three designs, one stable, one marginally unstable and one unstable, were used to demonstrate both an enhanced overall injector design process and an improved combustion stability assessment process. High-fidelity results from state-of-the-art computational fluid dynamics CFD simulations were used to substantially augment and improve the injector design methodology. The CFD results were used to inform and guide the overall injector design process. They were also used to upgrade selected empirical or low-dimensional quantities in the ROCket Combustor Interactive Design (ROCCID) stability assessment tool. Hot fire single element injector testing was used to verify both the overall injector designs and the stability assessments. Testing was conducted at the Air Force Research Laboratory and at Purdue University. Companion papers provide details of the overall injector design process, full- and sub-scale testing, ROCCID-based stability assessments and the CFD simulations.

Tucker, P. K.↗
NASA NTRS2015

The Effect of Resistance on Rocket Injector Acoustics

Combustion instability, where unsteady heat release couples with acoustic modes, has long been an area of concern in liquid rocket engines. Accurate modeling of the acoustic normal modes of the combustion chamber is important to understanding and preventing combustion instability. This study evaluates the effect of injector resistance on the mode shapes and complex eigen-frequencies of an injector/combustion chamber system by defining a high Mach-flow form of the convective wave equation (see Eq. 1) in COMSOL Multiphysics' Coefficient Form PDE Mathematics Module.

Morgan, C. J.↗
NASA NTRS2015

The Effect of Resistance on Rocket Injector Acoustics

Combustion instability, where unsteady heat release couples with acoustic modes, has long been an area of concern in liquid rocket engines. Accurate modeling of the acoustic normal modes of the combustion chamber is important to understanding and preventing combustion instability. The injector resistance can have a significant influence on the chamber normal mode shape, and hence on the system stability.

Morgan, C. J.↗
NASA NTRS2010

Performance and Stability Analyses of Rocket Thrust Chambers with Oxygen/Methane Propellants

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for future in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems developed by NASA, so limited test data and analysis results are available at this stage of early development. As part of activities for the Propulsion and Cryogenic Advanced Development (PCAD) project funded under the Exploration Technology Development Program, the NASA Marshall Space Flight Center (MSFC) has been evaluating capability to model combustion performance and stability for oxygen and methane propellants. This activity has been proceeding for about two years and this paper is a summary of results to date. Hot-fire test results of oxygen/methane propellant rocket engine combustion devices for the modeling investigations have come from several sources, including multi-element injector tests with gaseous methane from the 1980s, single element tests with gaseous methane funded through the Constellation University Institutes Program, and multi-element injector tests with both gaseous and liquid methane conducted at the NASA MSFC funded by PCAD. For the latter, test results of both impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interactive Design and Analysis code and the Coaxial Injector Combustion Model. Special effort was focused on how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied, improved or developed in the future. Low frequency combustion instability (chug) occurred, with frequencies ranging from 150 to 250 Hz, with several multi-element injectors with liquid/liquid propellants, and was modeled using techniques from Wenzel and Szuch. High-frequency combustion instability also occurred at the first tangential (1T) mode, at about 4500 Hz, with several multi-element injectors with liquid/liquid propellants. Analyses of the transverse mode instability were conducted by evaluating injector resonances and empirical methods developed by Hewitt.

Hulka, James R.↗
NASA NTRS2014

Mean Flow Augmented Acoustics in Rocket Systems

Combustion instability in solid rocket motors and liquid engines has long been a subject of concern. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process and gas dynamics. Recent advances in energy based modeling of combustion instabilities require accurate determination of acoustic frequencies and mode shapes. Of particular interest is the acoustic mean flow interactions within the converging section of a rocket nozzle, where gradients of pressure, density, and velocity become large. The expulsion of unsteady energy through the nozzle of a rocket is identified as the predominate source of acoustic damping for most rocket systems. Recently, an approach to address nozzle damping with mean flow effects was implemented by French [1]. This new approach extends the work originated by Sigman and Zinn [2] by solving the acoustic velocity potential equation (AVPE) formulated by perturbing the Euler equations [3]. The present study aims to implement the French model within the COMSOL Multiphysiscs framework and analyzes one of the author's presented test cases.

Fischbach, Sean↗
NASA NTRS2020

Combustion Dynamics Characteristics and Fuel Pressure Modulation Responses of a Three-Cup Third-Generation Swirl-Venturi Lean Direct Injection Combustion Concept

This paper presents the combustion dynamic data and fuel modulation response of a three-cup Lean Direct Injection combustor developed by Woodward, FST. The test was conducted at the NASA Glenn Research Center CE-5 flame tube test facility. The facility provided inlet air up to 922 K and pressure up to 19.0 bar. At the low-power configuration, the combustion noise was quiet. Large combustion pressure oscillations were observed with the High-power configuration at an off design condition, with low inlet air temperature and pressure conditions and a high equivalence ratio (about T3=600 K, P3 = 800 kPa, and ER =0.46). The noise amplitude was as high as 1.5 psi at around 220 Hz. As inlet air pressure and temperature increased, this combustion instability decreased. Fuel modulated signals were produced with the WASK fuel modulator located in the fuel line upstream of the center cup pilot fuel-air mixer. The amplitudes of the modulated signals detected in the combustor were low. Only less than 0.13% (0.06 psi) of the input energy was detected, and the signal amplitudes decreased as the modulated frequencies increased. Interaction between the modulated signals and the combustion noise varied with operating conditions. At a condition with low combustion noise around 150 hz, modulating a signal at around the same frequency would increase the combustion noise from 0.2 psi to as high as 0.6 psi, whereas at a condition with a high combustion instability around 250 hz, the modulated signal did not seem to have much effect on the combustion noise.

Fuel flexible combustor↗
NASA NTRS1959

Analysis of of Injection-Velocity Effects on Rocket Motor Dynamics and Stability

A concept of combustion time lag that includes dependency on injection velocity is introduced. The concept is used in the formulation of chamber transfer functions and in an analysis of low-frequency combustion instability. Theoretical frequency responses and stability boundaries are compared with those obtained when the injection-velocity effect on the time lag to be an important consideration, in the theory of chamber dynamics and combustion instability

Hurrell, Herbert G.↗
NASA NTRS2010

J-2X Gas Generator Development Testing at NASA Marshall Space Flight Center

NASA is developing a liquid oxygen/liquid hydrogen rocket engine for upper stage and trans-lunar applications of the Ares vehicles for the Constellation program. This engine, designated the J-2X, is a higher pressure, higher thrust variant of the Apollo-era J-2 engine. Development was contracted to Pratt & Whitney Rocketdyne in 2006. Over the past several years, two phases of testing have been completed on the development of the gas generator for the J-2X engine. The hardware has progressed through a variety of workhorse injector, chamber, and feed system configurations. Several of these configurations have resulted in combustion instability of the gas generator assembly. Development of the final configuration of workhorse hardware (which will ultimately be used to verify critical requirements on a component level) has required a balance between changes in the injector and chamber hardware in order to successfully mitigate the combustion instability without sacrificing other engine system requirements. This paper provides an overview of the two completed test series, performed at NASA s Marshall Space Flight Center. The requirements, facility setup, hardware configurations, and test series progression are detailed. Significant levels of analysis have been performed in order to provide design solutions to mitigate the combustion stability issues, and these are briefly covered. Also discussed are the results of analyses related to either anomalous readings or off-nominal testing throughout the two test series.

Reynolds, D. C.↗
NASA NTRS1973

Performance gains by using heated natural-gas fuel in an annular turbojet combustor

A full-scale annular turbojet combustor was tested with natural gas fuel heated from ambient temperature to 800 K (980 F). In all tests, heating the fuel improved combustion efficiency. Two sets of gaseous fuel nozzles were tested. Combustion instabilities occurred with one set of nozzles at two conditions: one where the efficiency approached 100 percent with the heated fuel; the other where the efficiency was very poor with the unheated fuel. The second set of nozzles exhibited no combustion instability. Altitude relight tests with the second set showed that relight was improved and was achievable at essentially the same condition as blowout when the fuel temperature was 800 K (980 F).

Marchionna, N. R.↗
NASA NTRS1979

Solid Rocket Motor Technology

These Conference Proceedings contain the 36 papers presented at the AGARD Propulsion and Energetics Panel 53rd Symposium on "Solid Rocket Motor Technology", which was held in Oslo, Norway, 2-5 April 1979. The questions and answers after the various presentations are included as well as the Round Table Discussion. The Symposium was organized into seven sessions: Survey Papers (2); Ignition, Extinction and Internal Ballistics (8); Burn Rate Modelling and Combustion of Metal (5); New Propellants (5); Combustion Instability (7); Heat Transfer and Materials (3); and Testing and Instrumentation (6).The Symposium provided a forum for discussion and exchange between research scientists ,propellant formulators and motor technologists. Though main emphasis was placed on transient processes like ignition and extinction and on combustion instability with the various related aspects there were a number of papers dealing with propellants, their characteristics and their development. Whether the symposium achieved its initial aim, i.e. furnishing a comprehensive survey of the technology available for solid propellant rocket motors and its further development capabilities, will be assessed in a Technical Evaluation Report, to be published as AGARD Advisory Report AR 151.

Advisory Group for Aerospace Research and Developm↗
NASA NTRS1975

Effect of Nozzle Nonlinearities upon Nonlinear Stability of Liquid Propellant Rocket Motors

A three dimensional, nonlinear nozzle admittance relation is developed by solving the wave equation describing finite amplitude oscillatory flow inside the subsonic portion of a choked, slowly convergent axisymmetric nozzle. This nonlinear nozzle admittance relation is then used as a boundary condition in the analysis of nonlinear combustion instability in a cylindrical liquid rocket combustor. In both nozzle and chamber analyses solutions are obtained using the Galerkin method with a series expansion consisting of the first tangential, second tangential, and first radial modes. Using Crocco's time lag model to describe the distributed unsteady combustion process, combustion instability calculations are presented for different values of the following parameters: (1) time lag, (2) interaction index, (3) steady-state Mach number at the nozzle entrance, and (4) chamber length-to-diameter ratio. In each case, limit cycle pressure amplitudes and waveforms are shown for both linear and nonlinear nozzle admittance conditions. These results show that when the amplitudes of the second tangential and first radial modes are considerably smaller than the amplitude of the first tangential mode the inclusion of nozzle nonlinearities has no significant effect on the limiting amplitude and pressure waveforms.

Padmanabhan, M. S.↗
NASA NTRS1978

Stability analysis of a liquid fuel annular combustion chamber

High frequency combustion instability problems in a liquid fuel annular combustion chamber are examined. A modified Galerkin method was used to produce a set of modal amplitude equations from the general nonlinear partial differential acoustic wave equation in order to analyze the problem of instability. From these modal amplitude equations, the two variable perturbation method was used to develop a set of approximate equations of a given order of magnitude. These equations were modeled to show the effects of velocity sensitive combustion instabilities by evaluating the effects of certain parameters in the given set of equations.

Mcdonald, G. H.↗
NASA NTRS1993

Inherent stability of central element coaxial liquid-liquid injectors

Most TRW liquid bi-propellant rocket engines built over the past thirty-plus years have employed a central element coaxial pintle injector and have operated with liquid/liquid propellant injection. This injector is a patented design exclusive to TRW and has unique features that make the rocket engine combustion characteristics different from those of other types of injector engine designs. Its many benefits include excellent combustion performance, efficient deep throttling, adaptability to low cost manufacturing, and high reliability. Approximately 200 pintle injector engines of various sizes and operating on a variety of propellants have been flown without a single inflight failure. An especially important feature of the pintle injector engine is its apparent inherent combustion stability. In over thirty years of development, testing, and production, TRW has never experienced combustion instability in any of its pintle injector engine designs. This has been true of engines operating over a range of thrust from 5 to 250,000 lbs. on earth-storable hypergolic propellants and a large number of smaller engines operating on a variety of propellants (21 combinations) in long duration-firing, pulsing (down to 2 msec), and deep throttling (as much as 19:1) modes. Operating chamber pressures have ranged from 10 to 3,500 psia. This record is particularly impressive given that typical TRW design practice does not consider combustion instability as an issue and no pintle engine has ever employed stability-enhancing features, such as baffles or acoustically resonant chambers. In spite of this, TRW engines have operated stably in regimes not possible with other types of injectors. Various physical explanations and combustion process models for this favorable stability characteristic have been postulated. However, a definitive study that unequivocally establishes the important stabilizing mechanisms still remains to be conducted.

Stoddard, Frank J.↗