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Temperature and density measurements in base region flow field of clustered rocket engine model using electron beam technique
An ignition quality tester was used to characterize the autoignition delay times of iso-octane. The experimental data were characterized between temperatures of 653 and 996 K, pressures of 1.0 and 1.5 MPa, and global equivalence ratios of 0.7 and 1.05. A clear negative temperature coefficient behavior was seen at both pressures in the experimental data. These data were used to characterize the effectiveness of three modeling methods: a single-zone homogeneous batch reactor, a multizone engine model, and a three-dimensional computational fluid dynamics (CFD) model. A detailed 874 species iso-octane ignition mechanism (Mehl, M.; Curran, H. J.; Pitz, W. J.; Westbrook, C. K.Chemical kinetic modeling of component mixtures relevant to gasoline. Proceedings of the European Combustion Meeting; Vienna, Austria, April 14-17, 2009) was reduced to 89 species for use in these models, and the predictions of the reduced mechanism were consistent with ignition delay times predicted by the detailed chemical mechanism across a broad range of temperatures, pressures, and equivalence ratios. The CFD model was also run without chemistry to characterize the extent of mixing of fuel and air in the chamber. The calculations predicted that the main part of the combustion chamber was fairly well-mixed at longer times (> ~30 ms), suggesting that the simpler models might be applicable in this quasi-homogeneous region. The multizone predictions, where the combustion chamber was divided into 20 zones of temperature and equivalence ratio, were quite close to the coupled CFD-kinetics results, but the calculation time was ~11 times faster than the coupled CFD-kinetics model. Although the coupled CFD-kinetics model captured the observed negative temperature coefficient behavior and pressure dependence, discrepancies remain between the predictions and the observed ignition time delays, suggesting improvements are still needed in the kinetic mechanism and/or the CFD model. This approach suggests a combined modeling approach, wherein the CFD calculations (without chemistry) can be used to examine the sensitivity of various model inputs to in-cylinder temperature and equivalence ratios. In conclusion, these values can be used as inputs to the multizone model to examine the impact on ignition delay. Additionally, the speed of the multizone model also makes it feasible to quickly test more detailed kinetic mechanisms for comparison to experimental data and sensitivity analysis.
SURF and SURFplus are high explosive reactive burn models for shock initiation and propagation of detonation waves. They are engineering models motivated by the ignition & growth concept of high spots and for SURFplus a second slow reaction for the energy release from carbon clustering. A key feature of the SURF model is that there is a partial decoupling between model parameters and detonation properties. This enables reduced sets of independent parameters to be calibrated sequentially for the initiation and propagation regimes. Here we focus on a methodology for tting the initiation parameters to Pop plot data based on 1-D simulations to compute a numerical Pop plot. In addition, the strategy for tting the remaining parameters for the propagation regime and failure diameter is discussed.
Current problems involving the structural integrity of propulsion systems, and proposed approaches to solving them, are reviewed. Areas investigated include the stall hammershock at the engine inlet, distortion-induced vibration in fan and compressor blading, modeling engine static structures with conical-shell finite elements, design and development of low-cost, self-contained bearing lubrication systems for turbine engines, and roller bearing slip and skidding damage. Individual items are announced in this issue.
The emission problem or source points were defined and new materials, hardware, or operational procedures were developed to exercise the trends defined by the data collected. The programs to reduce the emission output of aircraft powerplants were listed. Continued establishment of baseline emissions for various engine models, continued characterization of effect of production tolerances on emissions, carbureted engine development and flight tests, and cylinder cooling/fin design programs were several of the programs investigated.
Systematic scaling or dimensional analysis reveals that certain scales of geophysical fluid flows (such as stellar, ocean, and planetary atmosphere circulations) can be accurately modeled in the laboratory using a procedure which differs from conventional engineering modeling. Rather than building a model to obtain numbers for a specific design problem, the relative effects of the significant forces are systematically varied in an attempt to deepen understanding of the effects of these forces. Topics covered include: (1) modeling a large-scale planetary atmospheric flow in a rotating cylindrical annulus; (2) achieving a radial dielectric body force; (3) spherical geophysical fluid dynamics experiments for Spacelab flights; (4) measuring flow and temperature; and (5) the possible effect of rotational or precessional disturbances on the flow in the rotating spherical containers.
An advanced LOX/LH2 engine study for the use of NASA and vehicle prime contractors in developing concepts for manned missions to the Moon, Mars, and Phobos is documented. Parametric design data was obtained at five engine thrusts from 7.5K lbf to 50K lbf. Also, a separate task evaluated engine throttling over a 20:1 range and operation at a mixture ratio of 12 plus or minus 1 versus the 6 plus or minus 1 nominal. Cost data was also generated for DDT&E, first unit production, and factors in other life cycle costs. The major limitation of the study was lack of contact with vehicle prime contractors to resolve the issues in vehicle/engine interfaces. The baseline Aerojet dual propellant expander cycle was shown capable of meeting all performance requirements with an expected long operational life due to the high thermal margins. The basic engine design readily accommodated the 20:1 throttling requirement and operation up to a mixture ratio of 10 without change. By using platinum for baffled injector construction the increased thermal margin allowed operation up to mixture ratio 13. An initial engine modeling with an Aerojet transient simulation code (named MLETS) indicates stable engine operation with the baseline control system. A throttle ratio of 4 to 5 seconds from 10 percent to 100 percent thrust is also predicted. Performance predictions are 483.1 sec at 7.5K lbf, 487.3 sec at 20K lbf, and 485.2 sec at 50K lbf with a mixture ratio of 6 and an area ratio of 1200. Engine envelopes varied from 120 in. length/53 in. exit diameter at 7.5K lbf to 305 in. length/136 in. exit diameter at 50 K lbf. Packaging will be an important consideration. Continued work is recommended to include more vehicle prime contractor/engine contractor joint assessment of the interface issues.
The Linear Autoland Simulink model was created to be a modular test environment for testing of control system components in commercial aircraft. The input variables, physical laws, and referenced frames used are summarized. The state space theory underlying the model is surveyed and the location of the control actuators described. The equations used to realize the Dryden gust model to simulate winds and gusts are derived. A description of the pseudo-random number generation method used in the wind gust model is included. The longitudinal autopilot, lateral autopilot, automatic throttle autopilot, engine model and automatic trim devices are considered as subsystems. The experience in converting the Airlabs FORTRAN aircraft control system simulation to a graphical simulation tool (Matlab/Simulink) is described.
Experimental supersonic combustion research related to hypersonic airbreathing propulsion has been actively underway at NASA Langley Research Center (LaRC) since the mid-1960's. This research involved experimental investigations of fuel injection, mixing, and combustion in supersonic flows and numerous tests of scramjet engine flowpaths in LaRC test facilities simulating flight from Mach 4 to 8. Out of this research effort has come scramjet combustor design methodologies, ground test techniques, and data analysis procedures. These technologies have progressed steadily in support of the National Aero-Space Plane (NASP) program and the current Hyper-X flight demonstration program. During NASP nearly 2500 tests of 15 scramjet engine models were conducted in LaRC facilities. In addition, research supporting the engine flowpath design investigated ways to enhance mixing, improve and apply nonintrusive diagnostics, and address facility operation. Tests of scramjet combustor operation at conditions simulating hypersonic flight at Mach numbers up to 17 also have been performed in an expansion tube pulse facility. This paper presents a review of the LaRC experimental supersonic combustion research efforts since the late 1980's, during the NASP program, and into the Hyper-X Program.
The integration of flight control and propulsion control has been a much discussed topic, especially for emergencies where the engines may be able to help stabilize and safely land a damaged aircraft. Previous research has shown that for the engines to be effective as flight control actuators, the response time to throttle commands must be improved. Other work has developed control modes that accept a higher risk of engine failure in exchange for improved engine response during an emergency. In this effort, a nonlinear engine model (the Commercial Modular Aero-Propulsion System Simulation 40k) has been integrated with a nonlinear airframe model (the Generic Transport Model) in order to evaluate the use of enhanced-response engines as alternative yaw rate control effectors. Tests of disturbance rejection and command tracking were used to determine the impact of the engines on the aircraft's dynamical behavior. Three engine control enhancements that improve the response time of the engine were implemented and tested in the integrated simulation. The enhancements were shown to increase the engine s effectiveness as a yaw rate control effector when used in an automatic feedback loop. The improvement is highly dependent upon flight condition; the airframe behavior is markedly improved at low altitude, low speed conditions, and relatively unchanged at high altitude, high speed.
Computer simulation is an essential part of the design and development of jet engines for the aeropropulsion industry. Engineers concerned with calculating the flow in jet engine components, such as compressors and turbines, need simple engineering models that accurately describe the complex flow of air and gases and that allow them to quickly estimate loads, losses, temperatures, and other design parameters. In this ongoing collaborative project, advanced wavelet analysis techniques are being used to gain insight into the complex flow phenomena. These insights, which cannot be achieved by commonly used methods, are being used to develop innovative new flow models and to improve existing ones. Wavelet techniques are very suitable for analyzing the complex turbulent and transitional flows pervasive in jet engines. These flows are characterized by intermittency and a multitude of scales. Wavelet analysis results in information about these scales and their locations. The distribution of scales is equivalent to the frequency spectrum provided by commonly used Fourier analysis techniques; however, no localization information is provided by Fourier analysis. In addition, wavelet techniques allow conditional sampling analyses of the individual scales, which is not possible by Fourier methods.
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The data presented in this contribution were obtained in the NASA Langley 16-Foot Transonic Tunnel. Multiple test entries were completed and the results have been completely reported in five NASA reports. The objective of the initial investigation was to determine the effect of empennage (tail) interference on the drag characteristics of an axisymmetric model with a single engine fighter aft-end with convergent divergent nozzles. Two nozzle power settings, dry and maximum afterburning, were investigated. Several empennage arrangements and afterbody modifications were investigated during the initial investigation. Subsequent investigations were used to determine the effects of other model variables including tail incidence, tail span, and nozzle shape. For the final investigation, extensive surface pressure instrumentation was added to the model in order to develop and understanding of the flow interactions associated with afterbody/empennage integration and also to provide data for code validation. Extensive computational analysis has been conducted on the staggered empennage configuration at a Mach number of 0.6 utilizing a three-dimensional Navier Stokes code. Most of the investigations were conducted at Mach numbers from 0.60 to 1.20 and at ratios of jet total pressure to free stream static pressure (nozzle pressure ratio) from 0.1 (jet off) to 8.0. Some angle of attack variation was obtained at jet off conditions.
Complex physics and long computation time hinder the adoption of computer aided engineering models in the design of large-format battery cells and systems. A modular, efficient battery simulation model -- the multiscale multidomain (MSMD) model -- was previously introduced to aid the scale-up of Li-ion material and electrode designs to complete cell and pack designs, capturing electrochemical interplay with 3-D electronic current pathways and thermal response. Here, this paper enhances the computational efficiency of the MSMD model using a separation of time-scales principle to decompose model field variables. The decomposition provides a quasi-explicit linkage between the multiple length-scale domains and thus reduces time-consuming nested iteration when solving model equations across multiple domains. In addition to particle-, electrode- and cell-length scales treated in the previous work, the present formulation extends to bus bar- and multi-cell module-length scales. We provide example simulations for several variants of GH electrode-domain models.
Core-collapse supernova science is now entering an era where engine models are beginning to make both qualitative and quantitative predictions. Although the evidence in support of the convective engine for core-collapse supernova continues to grow, it is difficult to place quantitative constraints on this engine. Some studies have made specific predictions for the remnant distribution from the convective engine, but the results differ between different groups. Here we use a broad parameterization for the supernova engine to understand the differences between distinct studies. With this broader set of models, we place error bars on the remnant mass and basic yields from the uncertainties in the explosive engine. We fi nd that, even with only 3 progenitors and a narrow range of explosion energies, we can produce a wide range of remnant masses and nucleosynthetic yields.
In this paper, a heterogeneous propulsion system simulation method is presented. The method is based on the formulation of a cycle model of a gas turbine engine. The model includes the nonlinear characteristics of the engine components via use of empirical data. The potential to simulate the entire engine operation on a computer without the aid of data is demonstrated by numerically generating "performance maps" for a fan component using two flow models of varying fidelity. The suitability of the fan models were evaluated by comparing the computed performance with experimental data. A discussion of the potential benefits and/or difficulties in connecting simulations solutions of differing fidelity is given.
The CFR engine is the widely accepted platform to test standard Research Octane Number (RON) and Motored Octane Number (MON) for determining anti-knock characteristics of motor fuels. With increasing interest in engine downsizing, up-torquing, and alternative fuels for modern spark ignition (SI) engines, there is a need to better understand the conditions that fuels are subjected to in the CFR engine during octane rating. To take into account fuel properties, such as fuel heat of vaporization, laminar flame speed and auto-ignition chemistry; and understand their impacts on combustion knock, it is essential to estimate accurate cylinder conditions. In this study, the CFR F1/F2 engine was modeled using GT-Power with the Three Pressure Analysis (TPA) and the model was validated for different fuels and engine conditions. The finite element cylinder model was applied to better estimate heat transfer and cylinder wall temperatures of the cast iron combustion chamber (piston, cylinder, and head) of the CFR engine under continuous standard knocking operation. Uncertainty of unburned gas temperature and cylinder wall temperatures estimation was analyzed depending on uncertainty in experimentally measured and prescribed model input parameters. The model was finally used to estimate the IVC conditions (trapped mass, residual gas fraction, and temperature), as well as unburned gas temperature and cylinder wall temperatures for various fuels throughout the engine cycle.
An aircraft model that incorporates independently adjustable engine throttles and ailerons is employed to develop an adaptive control scheme in the presence of actuator failures. This model captures the key features of aircraft flight dynamics when in the engine differential mode. Based on this model an adaptive feedback control scheme for asymptotic state tracking is developed and applied to a transport aircraft model in the presence of two types of failures during operation, rudder failure and aileron failure. Simulation results are presented to demonstrate the adaptive failure compensation scheme.