Search NASA⌕ Search

SEARCH · Search NASA

Results for “COMBUSTION INSTABILITY”

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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Development of a computational testbed for numerical simulation of combustion instability

A synergistic hierarchy of analytical and computational fluid dynamic techniques is used to analyze three-dimensional combustion instabilities in liquid rocket engines. A mixed finite difference/spectral procedure is employed to study the effects of a distributed vaporization zone on standing and spinning instability modes within the chamber. Droplet atomization and vaporization are treated by a variety of classical models found in the literature. A multi-zone, linearized analytical solution is used to validate the accuracy of the numerical simulations at small amplitudes for a distributed vaporization region. This comparison indicates excellent amplitude and phase agreement under both stable and unstable operating conditions when amplitudes are small and proper grid resolution is used. As amplitudes get larger, expected nonlinearities are observed. The effect of liquid droplet temperature fluctuations was found to be of critical importance in driving the instabilities of the combustion chamber.

Grenda, Jeffrey↗
NASA NTRS1991

Computational fluid dynamic analysis of liquid rocket combustion instability

The paper presents a computational analysis of liquid rocket combustion instability. Consideration is given to both a fully nonlinear unsteady calculation as well as a new CFD-based linearized stability analysis. An analytical solution for the linear stability problem in a constant area combustion chamber with uniform mean flow is developed to verify the numerical analyses.

Venkateswaran, Sankaran↗
NASA NTRS1966

Nonlinear Aspects of Combustion Instability in Liquid Propellant Rocket Motors Sixth Yearly Progress Report, 1 Jun. 1965 - 31 May 1966

This is the sixth yearly report on the research at Princeton University on the problem of combustion instability in liquid-propellant rocket motors. These investigations have been sponsored by the National Aeronautics and Space Administration under NASA Contract NASr-217 and, in the past, under NASA Grant NsG-99-60. The nonlinear aspects of the problem have been emphasized in both the theoretical and experimental portions of the program. Presently, the theoretical research primarily involves studies of possible mechanisms or forcing functions of the instabilities, the nonlinear gasdynamic oscillations associated with instability in rocket motors, the effect of the exhaust nozzle on the stability, and the effect of damping devices on the stability. In conjunction with the theoretical program, an experimental program of an equally broad range is being conducted. This program includes basic experiments both with rocket motors and "cold" chambers which have been designed to study various combustion processes in order to provide understanding which would be useful in the construction of a theoretical model. It also includes studies of the effects of exhaust nozzles and damping devices upon the instability as well as controlled testing of liquid rocket motors with parametric variations intended to determine the factors which strongly influence the incidence of nonlinear combustion instability.

Combustion instability↗
NASA NTRS2004

Combustion Instability Analysis and the Effects of Drop Size on Acoustic Driving Rocket Flow

High frequency combustion instability, the most destructive kind, is generally solved on a per engine basis. The instability often is the result of compounding acoustic oscillations, usually from the propellant combustion itself. To counteract the instability the chamber geometry can be changed and/or the method of propellant injection can be altered. This experiment will alter the chamber dimensions slightly; using a cylindrical shape of constant diameter and the length will be varied from six to twelve inches in three-inch increments. The main flowfield will be the products of a high OF hydrogen/oxygen flow. The liquid fuel will be injected into this flowfield using a modulated injector. It will allow for varied droplet size, feed rate, spray pattern, and location for the mixture within the chamber. The response will be deduced from the chamber pressure oscillations.

Harper, Brent↗
NASA NTRS2005

Optical Sensing of Combustion Instabilities in Gas Turbines

In a continuing program of research and development, a system has been demonstrated that makes high-speed measurements of thermal infrared radiance from gas-turbine engine exhaust streams. When a gas-turbine engine is operated under conditions that minimize the emission of pollutants, there is a risk of crossing the boundary from stable to unstable combustion. Combustion instability can lead to engine damage and even catastrophic failure. Sensor systems of the type under development could provide valuable data during the development testing of gas-turbine engines or of engine components. A system of the type under development makes high-speed measurements of thermal infrared radiance from the engine exhaust stream. The sensors of this system can be mounted outside the engine, which eliminates the need for engine case penetrations typical with other engine dynamics monitors. This is an important advantage in that turbine-engine manufacturers consider such penetrations to be very undesirable. A prototype infrared sensor system has been built and demonstrated on a turbine engine. This system includes rugged and inexpensive near-infrared sensors and filters that select wavelengths of infrared radiation for high sensitivity. In experiments, low-frequency signatures were consistently observed in the detector outputs. Under some conditions, the signatures also included frequency components having one or two radiance cycles per engine revolution. Although it has yet to be verified, it is thought that the low-frequency signatures may be associated with bulk-mode combustion instabilities or flow instabilities in the compressor section of the engine, while the engine- revolution-related signatures may be indicative of mechanical problems in the engine. The system also demonstrated the ability to detect transient high-radiance events. These events indicate hot spots in the exhaust stream and were found to increase in frequency during engine acceleration.

Markham, James R.↗
NASA NTRS1990

Navier-Stokes Entropy Controlled Combustion Instability Analysis for Liquid Propellants

Navier-Stokes solutions are used to calculate oscillatory components of pressure, velocity, and density, which in turn provide necessary data to compute energy growth factors to determine combustion instability. It is shown that wave instabilities are associated with changes in entropy and the space and time averages of oscillatory components of pressure, velocity and density, together with the mean flow field in the energy equation. Compressible laminar and turbulent flows and reacting flows with hydrogen/oxygen combustion are considered. The SSME combustion/thrust chamber is used for illustration of the theory. The analysis shows that the increase of mean pressure and disturbances consistently results in the increase of instability. It is shown that adequate combustion instability analysis requires at least third order nonlinearity in energy growth or decay.

Chung, T. J.↗
NASA NTRS1992

Multi-dimensional analysis of combustion instabilities in liquid rocket motors

A three-dimensional analysis of combustion instabilities in liquid rocket engines is presented based on a mixed finite difference/spectral solution methodology for the gas phase and a discrete droplet tracking formulation for the liquid phase. Vaporization is treated by a simplified model based on an infinite thermal conductivitiy assumption for spherical liquid droplets of fuel in a convective environment undergoing transient heating. A simple two parameter phenomenological combustion response model is employed for validation of the results in the small amplitude regime. The computational procedure is demonstrated to capture the phenomena of wave propagation within the combustion chamber accurately. Results demonstrate excellent amplitude and phase agreement with analytical solutions for properly selected grid resolutions under both stable and unstable operating conditions. Computations utilizing the simplified droplet model demonstrate stable response to arbitrary pulsing. This is possibly due to the assumption of uniform droplet temperature which removes the thermal inertia time-lag response of the vaporization process. The mixed-character scheme is sufficiently efficient to allow solutions on workstations at a modest increase in computational time over that required for two-dimensional solutions.

Grenda, Jeffrey M.↗
NASA NTRS1993

Numerical parametric studies of spray combustion instability

A coupled numerical algorithm has been developed for studies of combustion instabilities in spray-driven liquid rocket engines. The model couples gas and liquid phase physics using the method of fractional steps. Also introduced is a novel, efficient methodology for accounting for spray formation through direct solution of liquid phase equations. Preliminary parametric studies show marked sensitivity of spray penetration and geometry to droplet diameter, considerations of liquid core, and acoustic interactions. Less sensitivity was shown to the combustion model type although more rigorous (multi-step) formulations may be needed for the differences to become apparent.

Pindera, M. Z.↗
NASA NTRS1998

Elimination of High-Frequency Combustion Instability in the Fastrac Engine Thrust Chamber

NASA's Marshall Space Flight Center(MSFC) has been tasked with developing a 60,000 pound thrust, pump-fed, LOX/RP-1 engine under the Advanced Space Transportation Program(ASTP). This government-led design has been designated the Fastrac engine. The X-34 vehicle will use the Fastrac engine as the main propulsion system. The X-34 will be a suborbital vehicle developed by the Orbital Sciences Corporation. The X-34 vehicle will be launched from an L-1011 airliner. After launch, the X-34 vehicle will be able to climb to altitudes up to 250,000 feet and reach speeds up to Mach 8, over a mission range of 500 miles. The overall length, wingspan, and gross takeoff weight of the X-34 vehicle are 58.3 feet, 27.7 feet and 45,000 pounds, respectively. This report summarizes the plan of achieving a Fastrac thrust chamber assembly(TCA) stable bomb test that meets the JANNAF standards, the Fastrac TCA design, and the combustion instabilities exhibited by the Fastrac TCA during testing at MSFC's test stand 116 as determined from high-frequency fluctuating pressure measurements. This report also summarizes the characterization of the combustion instabilities from the pressure measurements and the steps taken to eliminate the instabilities.

Rocker, Marvin↗