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Fluid dynamic pancake model and droplet evaporation and combustion analysis related to nonlinear combustion instability in liquid propellant rocket motors
Combustion instability, which results from a coupling of the combustion process and the fluid dynamics of the engine system, was investigated. The design of devices which reduce coupling (combustion chamber baffles) and devices which increase damping (acoustic absorbers) are described. Included in the discussion are design criteria and recommended practices, structural and mechanical design, thermal control, baffle geometry, baffle/engine interactions, acoustic damping analysis, and absorber configurations.
Fluid dynamic model and diffusion-controlled flame analysis data on nonlinear combustion instability in liquid propellant rocket engines
A computational fluid dynamics (CFD) code is being developed to compute the mixing and combustion of hydrogen fuel in the turbulent flow fields of supersonic combustion ramjets (scramjet). The code solves the three-dimensional Reynolds time-averaged complete Navier-Stokes equations including transport equations for a four species, two reaction, global finite rate chemistry model. The code was applied to the case of transverse injection of hydrogen from a sonic circular orifice into a supersonic airstream. The equations were numerically integrated using MacCormack's explicit method, and the algebraic eddy viscosity model of Baldwin-Lomax was used to model the turbulence. In the species transport and energy equations, diffusion coefficients based on Fick's Law and an assumption of unit Lewis number were applied. Computed features of the three-dimensional flow field are depicted by static pressure, static temperature, mass fraction of species, and velocity vectors. For engineering interest, mixing and combustion parameters were examined to assess the effect of injector diameter, injected fuel pressure, fuel-air ratio, and spacing of fuel injectors. The objective of the present paper is to demonstrate the capability of the present three-dimensional spatially elliptic, CFD code for turbulent, reacting flow. Application of the code to specific supersonic combustion configurations is planned.
Injector design is a critical part of the development of a rocket Thrust Chamber Assembly (TCA). Proper detailed injector design can maximize propulsion efficiency while minimizing the potential for failures in the combustion chamber. Traditional design and analysis methods for hydrocarbon-fuel injector elements are based heavily on empirical data and models developed from heritage hardware tests. Using this limited set of data produces challenges when trying to design a new propulsion system where the operating conditions may greatly differ from heritage applications. Time-accurate, Three-Dimensional (3-D) Computational Fluid Dynamics (CFD) modeling of combusting flows inside of injectors has long been a goal of the fluid analysis group at Marshall Space Flight Center (MSFC) and the larger CFD modeling community. CFD simulation can provide insight into the design and function of an injector that cannot be obtained easily through testing or empirical comparisons to existing hardware. However, the traditional finite-rate chemistry modeling approach utilized to simulate combusting flows for complex fuels, such as Rocket Propellant-2 (RP-2), is prohibitively expensive and time consuming even with a large amount of computational resources. MSFC has been working, in partnership with Streamline Numerics, Inc., to develop a computationally efficient, flamelet-based approach for modeling complex combusting flow applications. In this work, a flamelet modeling approach is used to simulate time-accurate, 3-D, combusting flow inside a single Gas Centered Swirl Coaxial (GCSC) injector using the flow solver, Loci-STREAM. CFD simulations were performed for several different injector geometries. Results of the CFD analysis helped guide the design of the injector from an initial concept to a tested prototype. The results of the CFD analysis are compared to data gathered from several hot-fire, single element injector tests performed in the Air Force Research Lab EC-1 test facility located at Edwards Air Force Base.
Combustion technology and its environmental concerns are being considered with increasing attention, not only for global-scale effects, but also for toxicological implications, particularly in the lift conditions of traffic-congested areas and industrial sites. Majority combustion by-products (CO, NO(sub x)) and unburned hydrocarbons (HC), are already subject to increasingly severe regulations; however other, non-regulated minority species, mainly soot and heavy aromatic molecules, involve higher health risks, as they are suspected to be agents of serious pathologies and even mutagenic effects. This is but one of the reasons why much research work is being carried out worldwide on the physical properties of these substances. Correspondingly, the need arises to detect their presence in urban environments, with as high a sensitivity as is required by their low concentrations, proper time- and space-resolutions, and 'real-time' capabilities. Lidar techniques are excellent candidates to this purpose, although severe constraints limit their applicability, eye-safety problems and aerosol Mie scattering uncertainties above all. At CNR's Istituto Motori in Napels, a Lidar-like diagnostic system is being developed, aimed primarily at monitoring the dynamic behavior of internal combustion engines, particularly diesel exhausts, and at exploring the feasibility of a so-called 'Downtown Lidar'.
The gas turbine industry is directing particular attention to very low NOx combustors, whether for aircraft or land based CCGT systems. These low NOx combustors frequently use liquid fuels or natural gas burning under very lean premixed conditions with air or under rich-lean conditions, although only the first case is studied here. In land based systems, diluted steam or nitrogen are sometimes injected into the combustion chambers to reduce flame temperature. The NOx emissions from such systems are the product of three chemical mechanisms which are interrelated: the hydrocarbon prompt NO, the thermal NO (extended Zeldovich mechanism), and the nitrous oxide route to NO. Formation of NO2 from NO also occurs, as well as emission of carbon monoxide and unburnt hydrocarbons. When the fuel-oxidant proportion decreases towards leaner conditions, flame temperatures are lowered, resulting in the total NO being reduced and the thermal-NO contribution greatly diminished to the benefit of the remaining two mechanisms of NO formation. While knowledge of the elementary reactions and their chemical kinetics concerning methane and simple hydrocarbons combustion has existed for a number of years, its use for computer modeling is limited to simple flow dynamics configurations. Nevertheless, understanding of such combusting flows under a wide range of experimental conditions allows for analogies or speculations with more complex actual systems. Such understanding can be achieved by means of one dimensional laminar premixed flame modeling, with a full chemical mechanism which incorporates the three routes of NO formation. Complementary to this understanding is the modeling of the actual combustion system using a full description of the fluid dynamics coupled with a reduced chemical scheme, which is then compared against the first model. The objective of this investigation is to evaluate the relative importance of the three mechanisms of NO formation in lean premixed methane-air combustion with increasing pressure using the one dimensional plug flow package PREMIX, and to test the validity of a three dimensional model with a global chemical mechanism against the one dimensional model in the atmospheric pressure case. Methane is chosen because it is the only mechanism which is reasonably well known and is a good guide to the behavior of other hydrocarbons. The mixture ratio chosen is richer than that in lean gas turbines, but the combustion of this mixture with the low preheat gives realistic gas turbine final flame temperatures. Conditions of NO2 formation are also analyzed in the one dimensional model and results extrapolated to the case of gas turbines.
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.
The objective of the present flight experiment definition study is to investigate the effects of slow forced convective flows on the dynamics of isolated single droplet combustion and is designed to complement the quiescent, microgravity droplet combustion experiments (DCE-1 and DCE-2) of Williams and Dryer. The fuels selected for this study are the same as those of DCE, namely, a sooting alkane fuel (heptane) and a non-sooting alcohol (methanol), and imposed flow rates are chosen between 0 and 20 cm/s with varying ambient oxygen concentrations and pressures. Within this velocity range, both accelerating and decelerating flow effects will also be investigated. Two different approaches to generate the forced flow are currently under development in ground-based facilities; the first is a flow tunnel concept where the forced flow is imposed against a stationary droplet, and in the second a tethered droplet is translated at a specified velocity in a quiescent ambient medium. Depending upon the engineering feasibility a selection will be made between these two approaches so that the experiment can be accommodated in the Multiple Droplet Combustion Apparatus (MDCA) currently being designed for the International Space Station. Recently, we have finished designing and fabricating the experimental rigs using both the above mentioned concepts. The flow tunnel concept is implemented in a 2.2 second drop tower rig. Preliminary experiments have been carried out using heptane and methanol in air at atmospheric pressure. The translating droplet apparatus is scheduled to be tested in the 5 second drop facility in the near future. This report presents some of the experimental results obtained for heptane.
The steady state thermodynamic cycle balance of the single preburner staged combustion engine, coupled with dynamic transient analyses, dictated in detail the location and requirements for each valve defined in this volume. Valve configuration selections were influenced by overall engine and vehicle system weight and failure mode determinations. Modulating valve actuators are external to the valve and are line replaceable. Development and satisfactory demonstration of a high pressure dynamic shaft seal has made this configuration practical. Pneumatic motor driven actuators that use engine pumped hydrogen gas as the working fluid are used. The helium control system is proposed as a module containing a cluster of solenoid actuated valves. The separable couplings and flanges are designed to assure minimum leakage with minimum coupling weight. The deflection of the seal surface in the flange is defined by finite element analysis that has been confirmed with test data. The seal design proposed has passed preliminary pressure cycling and thermal cycling tests.
The following areas of flight science are discussed in detail; (1) inviscid flow, (2) viscous flow, (3) aircraft aerodynamics, (4) fluid mechanics, (5) propulsion and combustion, and (6) flight dynamics and control.
The report presents fifteen papers from a workshop on microgravity science and applications held at the Jet Propulsion Laboratory in Pasadena, California, on December 3 to 4, 1984. The workshop and panel were formed by the Solid State Sciences Committee of the Board on Physics and Astronomy of the National Research Council in response to a request from the Office of Science and Technology Policy. The goal was to review the microgravity science and applications (MSA) program of NASA and to evaluate the quality of the program. The topics for the papers are metals and alloys, electronic materials, ceramics and glasses, biotechnology, combustion science, and fluid dynamics.
The KIVA-2 code, originally developed to solve computational fluid dynamics problems in internal combustion engines, has been developed to solve rocket propulsion type flows. The objective of the work was to develop a code such that both liquid and solid particle motion could be simulated for arbitrary geometry and high speed as well as low speed reacting flows. Modification to the original code include: incorporating independently specific supersonic and subsonic inflows and outflows; symmetric as well as periodic boundary conditions; and the capability to use generalized single or multi-specie thermodynamic data and transport coefficients allowing the user to specify arbitrary wall temperature/heat flux distributions. This code has been shown to successfully solve rocket propulsion flows as well as flows with entrained particles for several different rocket nozzles.
The measurement of the temporal distribution of fuel in gas turbine combustors is important in considering pollution, combustion efficiency and combustor dynamics and acoustics. Much of the previous work in measuring fuel distributions in gas turbine combustors has focused on the spatial aspect of the distribution. The temporal aspect however, has often been overlooked, even though it is just as important. In part, this is due to the challenges of applying real-time diagnostic techniques in a high pressure and high temperature environment. A simple and low-cost instrument that non-intrusively measures the real-time fuel-to-air ratio (FAR) in a gas turbine combustor has been developed. The device uses a dual wavelength laser absorption technique to measure the concentration of most hydrocarbon fuels such as jet fuel, methane, propane, etc. The device can be configured to use fiber optics to measure the local FAR inside a high pressure test rig without the need for windows. Alternatively, the device can readily be used in test rigs that have existing windows without modifications. An initial application of this instrument was to obtain time-resolved measurements of the FAR in the premixer of a lean premixed prevaporized (LPP) combustor at inlet air pressures and temperatures as high as 17 atm at 800 K, with liquid JP-8 as the fuel. Results will be presented that quantitatively show the transient nature of the local FAR inside a LPP gas turbine combustor at actual operating conditions. The high speed (kHz) time resolution of this device, combined with a rugged fiber optic delivery system, should enable the realization of a flight capable active-feedback and control system for the abatement of noise and pollutant emissions in the future. Other applications that require an in-situ and time-resolved measurement of fuel vapor concentrations should also find this device to be of use.
A one-formula surrogate fuel formulation and its quasi-global combustion kinetics model are developed to support the design of injectors and thrust chambers of kerosene-fueled rocket engines. This surrogate fuel model depicts a fuel blend that properly represents the general physical and chemical properties of kerosene. The accompanying gaseous-phase thermodynamics of the surrogate fuel is anchored with the heat of formation of kerosene and verified by comparing a series of one-dimensional rocket thrust chamber calculations. The quasi-global combustion kinetics model consists of several global steps for parent fuel decomposition, soot formation, and soot oxidation and a detailed wet-CO mechanism to complete the combustion process. The final thermophysics formulations are incorporated with a computational fluid dynamics model for prediction of the combustion efficiency of an unielement, tripropellant combustor and the radiation of a kerosene-fueled thruster plume. The model predictions agreed reasonably well with those of the tests.
Proper control of the numerical dissipation/filter to accurately resolve all relevant multiscales of complex flow problems while still maintaining nonlinear stability and efficiency for long-time numerical integrations poses a great challenge to the design of numerical methods. The required type and amount of numerical dissipation/filter are not only physical problem dependent, but also vary from one flow region to another. This is particularly true for unsteady high-speed shock/shear/boundary-layer/turbulence/acoustics interactions and/or combustion problems since the dynamics of the nonlinear effect of these flows are not well-understood. Even with extensive grid refinement, it is of paramount importance to have proper control on the type and amount of numerical dissipation/filter in regions where it is needed.
A computational method for the analysis of longitudinal-mode liquid rocket combustion instability has been developed based on the unsteady, quasi-one-dimensional Euler equations where the combustion process source terms were introduced through the incorporation of a two-zone, linearized representation: (1) A two-parameter collapsed combustion zone at the injector face, and (2) a two-parameter distributed combustion zone based on a Lagrangian treatment of the propellant spray. The unsteady Euler equations in inhomogeneous form retain full hyperbolicity and are integrated implicitly in time using second-order, high-resolution, characteristic-based, flux-differencing spatial discretization with Roe-averaging of the Jacobian matrix. This method was initially validated against an analytical solution for nonreacting, isentropic duct acoustics with specified admittances at the inflow and outflow boundaries. For small amplitude perturbations, numerical predictions for the amplification coefficient and oscillation period were found to compare favorably with predictions from linearized small-disturbance theory as long as the grid exceeded a critical density (100 nodes/wavelength). The numerical methodology was then exercised on a generic combustor configuration using both collapsed and distributed combustion zone models with a short nozzle admittance approximation for the outflow boundary. In these cases, the response parameters were varied to determine stability limits defining resonant coupling onset.