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At least 217 records · Page 12

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NASA NTRS1991

Numerical simulation of fluid flow around a scramaccelerator projectile

Numerical simulations of the fluid motion and temperature distribution around a 'scramaccelerator' projectile are obtained for Mach numbers in the 5-10 range. A finite element method is used to solve the equations of motion for inviscid and viscous two-dimensional or axisymmetric compressible flow. The time-dependent equations are solved explicitly, using bilinear isoparametric quadrilateral elements, mass lumping, and a shock-capturing Petrov-Galerkin formulation. Computed results indicate that maintaining on-design performance for controlling and stabilizing oblique detonation waves is critically dependent on projectile shape and Mach number.

Pepper, Darrell W.↗

NASA NTRS1991

Navier-Stokes simulation of the supersonic combustion flowfield in a ram accelerator

A computational study of the ram accelerator, a ramjet-in-tube device for accelerating projectiles to ultrahigh velocities, is presented. The analysis is performed using a fully implicit TVD scheme that efficiently solves the Reynolds-averaged Navier-Stokes equations and the species continuity equations associated with a finite rate combustion model. The present results indicate that viscous effects are of primary importance in all the cases studied, shock-induced combustion always started in the boundary layer. The effects of Mach number, mixture composition, pressure and tubulence are investigated for various configurations. Two types of combustion processes, one stable and the other unstable, were observed depending on the inflow conditions. The possibility of stabilizing the detonation wave by means of a backward facing step is also investigated. Two numerical techniques were tested: vector extrapolation, to accelerate convergence, and a diagonal formulation that eliminates the expense of inverting large block matrices which arise in chemically reacting flows.

Yungster, Shaye↗

NASA NTRS1992

Study of shock-induced combustion using an implicit TVD scheme

The supersonic combustion flowfields associated with various hypersonic propulsion systems, such as the ram accelerator, the oblique detonation wave engine, and the scramjet, are being investigated using a new computational fluid dynamics (CFD) code. The code solves the fully coupled Reynolds-averaged Navier-Stokes equations and species continuity equations in an efficient manner. It employs an iterative method and a second order differencing scheme to improve computational efficiency. The code is currently being applied to study shock wave/boundary layer interactions in premixed combustible gases, and to investigate the ram accelerator concept. Results obtained for a ram accelerator configuration indicate a new combustion mechanism in which a shock wave induces combustion in the boundary layer, which then propagates outward and downstream. The combustion process creates a high pressure region over the back of the projectile resulting in a net positive thrust forward.

Yungster, Shayne↗

NASA NTRS1990

NAS (Numerical Aerodynamic Simulation Program) technical summaries, March 1989 - February 1990

Given here are selected scientific results from the Numerical Aerodynamic Simulation (NAS) Program's third year of operation. During this year, the scientific community was given access to a Cray-2 and a Cray Y-MP supercomputer. Topics covered include flow field analysis of fighter wing configurations, large-scale ocean modeling, the Space Shuttle flow field, advanced computational fluid dynamics (CFD) codes for rotary-wing airloads and performance prediction, turbulence modeling of separated flows, airloads and acoustics of rotorcraft, vortex-induced nonlinearities on submarines, and standing oblique detonation waves.

Source record↗

NASA NTRS1992

Parametric and time resolved studies of autoignition and flameholding in a clean-air supersonic combustor

A model scramjet combustor was tested using transverse sonic injection of gaseous hydrogen behind a rearward facing step. It was found that ignition in the constant area duct initiates well downstream and propagates upstream with velocities ranging from 300 to nearly 1,000 m/s in the laboratory frame. The downstream ignition develops into a detonation wave prior to reaching the region of observation. A number of parameters including fuel flow rate, air stagnation temperature and pressure, and the operating conditions of the downstream air ejector affected ignition conditions.

Whitehurst, R. B.↗

NASA NTRS1992

Experimental Investigation of Nozzle/Plume Aerodynamics at Hypersonic Speeds

The work performed by D. W. Bogdanoff and J.-L. Cambier during the period of 1 Feb. - 31 Oct. 1992 is presented. The following topics are discussed: (1) improvement in the operation of the facility; (2) the wedge model; (3) calibration of the new test section; (4) combustor model; (5) hydrogen fuel system for combustor model; (6) three inch calibration/development tunnel; (7) shock tunnel unsteady flow; (8) pulse detonation wave engine; (9) DCAF flow simulation; (10) high temperature shock layer simulation; and (11) the one dimensional Godunov CFD code.

Heinemann, K.↗

NASA NTRS1992

Numerical simulation of shock-induced combustion/detonation

A numerical study is conducted to simulate the shock-induced combustion in premixed H2-air mixtures. Two types of bodies, blunt (spherical projectile) and sharp (wedge), are considered in the study. A nine-species, 18-step finite-rate H2-air chemical reaction mechanism coupled with the Navier-Stokes equations is solved. The flow field over the blunt body is found to be unsteady, when the projectile velocity is same as the Chapman-Jouget velocity of the mixture. The unsteadiness is caused by the periodic instabilities originating in the stagnation zone. Numerical results show good qualitative agreement with the ballistic range shadowgraph. In addition, the frequency of oscillations, determined by using the Fourier power spectrum, is found to be in good agreement with the experiment. The flow field over the wedge is found to be stable for the conditions considered in this study. The oblique detonation wave structure is investigated and the important flow features are discussed.

Singh, D. J.↗

NASA NTRS1948

Knocking Combustion Observed in a Spark-Ignition Engine with Simultaneous Direct and Schlieren High-Speed Motion Pictures and Pressure Records

Simultaneous direct and Schlieren photographs at 40,000 frames per second and correlated pressure records were taken of knocking combustion in a special spark-ignition engine to ascertain the intensity of certain end-zone reactions previously noted from Schlieren photography alone. A violent propagated homogeneous autoignition, or a similar phenomenon, previously observed, was again observed. The pressure records show autoignition of varying violence before the passage of a probable detonation wave. Extensive autoignition without occurrence of gas vibrations was seen in one explosion.

Osterstrom, Gordon E↗

NASA NTRS1994

Experimental investigation of nozzle/plume aerodynamics at hypersonic speeds

Much of the work involved the Ames 16-Inch Shock Tunnel facility. The facility was reactivated and upgraded, a data acquisition system was configured and upgraded several times, several facility calibrations were performed and test entries with a wedge model with hydrogen injection and a full scramjet combustor model, with hydrogen injection, were performed. Extensive CFD modeling of the flow in the facility was done. This includes modeling of the unsteady flow in the driver and driven tubes and steady flow modeling of the nozzle flow. Other modeling efforts include simulations of non-equilibrium flows and turbulence, plasmas, light gas guns and the use of non-ideal gas equations of state. New experimental techniques to improve the performance of gas guns, shock tubes and tunnels and scramjet combustors were conceived and studied computationally. Ways to improve scramjet engine performance using steady and pulsed detonation waves were also studied computationally. A number of studies were performed on the operation of the ram accelerator, including investigations of in-tube gasdynamic heating and the use of high explosives to raise the velocity capability of the device.

Bogdanoff, David W.↗

NASA NTRS1996

Inlet and Propulsion Integration of Scram Propelled Vehicles

The material to be presented in these two lectures begins with cycle considerations of the turbojet engine combined with a ramjet engine to provide thrust over the range of Mach 0 to 5. We will then examine in some detail the aerodynamic behavior that occurs in the inlet operating near the peak speed. Following that, we shall view a numerical simulation through a baseline scramjet engine, starting at the entrance to the inlet, proceeding into the combustor and through the nozzle. In the next segment, we examine a combined rocket and ramjet propulsion system. Analysis and test results will be examined with a view toward evaluation of the concept as a practical device. Two other inlets will then be reviewed: a Mach 12 inlet and a Mach 18 configuration. Finally, we close our lectures with a discussion of the Detonation Wave engine, and inspect the physical and chemical behavior obtained from numerical simulation. A few final remarks will be made regarding the application of CFD for hypersonic propulsion components.

Povinelli, Louis A.↗

NASA NTRS2000

PDE Nozzle Optimization Using a Genetic Algorithm

Genetic algorithms, which simulate evolution in natural systems, have been used to find solutions to optimization problems that seem intractable to standard approaches. In this study, the feasibility of using a GA to find an optimum, fixed profile nozzle for a pulse detonation engine (PDE) is demonstrated. The objective was to maximize impulse during the detonation wave passage and blow-down phases of operation. Impulse of each profile variant was obtained by using the CFD code Mozart/2.0 to simulate the transient flow. After 7 generations, the method has identified a nozzle profile that certainly is a candidate for optimum solution. The constraints on the generality of this possible solution remain to be clarified.

Billings, Dana↗

NASA NTRS2001

Integrated Pulse Detonation Propulsion and Magnetohydrodynamic Power

The prospects for realizing an integrated pulse detonation propulsion and magnetohydrodynamic (MHD) power system are examined. First, energy requirements for direct detonation initiation of various fuel-oxygen and fuel-air mixtures are deduced from available experimental data and theoretical models. Second, the pumping power requirements for effective chamber scavenging are examined through the introduction of a scavenging ratio parameter and a scavenging efficiency parameter. A series of laboratory experiments were carried out to investigate the basic engineering performance characteristics of a pulse detonation-driven MHD electric power generator. In these experiments, stoichiometric oxy-acetylene mixtures seeded with a cesium hydroxide/methanol spray were detonated at atmospheric pressure in a 1-m-long tube having an i.d. of 2.54 cm. Experiments with a plasma diagnostic channel attached to the end of the tube confirmed the attainment of detonation conditions (p2/p1 approximately 34 and D approximately 2,400 m/sec) and enabled the direct measurement of current density and electrical conductivity (approximately = 6 S/m) behind the detonation wave front, In a second set of experiments, a 30-cm-long continuous electrode Faraday channel, having a height of 2.54 cm and a width of 2 cm, was attached to the end of the tube using an area transition duct. The Faraday channel was inserted in applied magnetic fields of 0.6 and 0.95 T, and the electrodes were connected to an active loading circuit to characterize power extraction dependence on load impedance while also simulating higher effective magnetic induction. The experiments indicated peak power extraction at a load impedance between 5 and 10 Omega. The measured power density was in reasonable agreement with a simple electrodynamic model incorporating a correction for near-electrode potential losses. The time-resolved thrust characteristics of the system were also measured, and it was found that the NM interaction exerted a negligible influence on system thrust and that the measured I(sub sp) of the system (200 see) exceeded that computed for an equivalent nozzleless rocket (120 see).

Litchford, Ron J.↗

NASA NTRS2001

Integrated Pulse Detonation Propulsion and Magnetohydrodynamic Power

The prospects for realizing an integrated pulse detonation propulsion and magnetohydrodynamic (MHD) power system are examined. First, energy requirements for direct detonation initiation of various fuel-oxygen and fuel-air mixtures are deduced from available experimental data and theoretical models. Second, the pumping power requirements for effective chamber scavenging are examined through the introduction of a scavenging ratio parameter and a scavenging efficiency parameter. A series of laboratory experiments were carried out to investigate the basic engineering performance characteristics of a pulse detonation-driven MHD electric power generator. In these experiments, stoichiometric oxy-acetylene mixtures seeded with a cesium hydroxide/methanol spray were detonated at atmospheric pressure in a 1-m-long tube having an i.d. of 2.54 cm. Experiments with a plasma diagnostic channel attached to the end of the tube confirmed the attainment of detonation conditions (p(sub 2)/p(sub 1) approx. 34 and D approx. 2,400 m/sec) and enabled the direct measurement of current density and electrical conductivity (=6 S/m) behind the detonation wave front. In a second set of experiments, a 30-cm-long continuous electrode Faraday channel, having a height of 2.54 cm and a width of 2 cm, was attached to the end of the tube using an area transition duct. The Faraday channel was inserted in applied magnetic fields of 0.6 and 0.95 T. and the electrodes were connected to an active loading circuit to characterize power extraction dependence on load impedance while also simulating higher effective magnetic induction. The experiments indicated peak power extraction at a load impedance between 5 and 10 Ohm. The measured power density was in reasonable agreement with a simple electrodynamic model incorporating a correction for near-electrode potential losses. The time-resolved thrust characteristics of the system were also measured, and it was found that the MHD interaction exerted a negligible influence on system thrust and that the measured I(sub sp) of the system (200 sec) exceeded that computed for an equivalent nozzleless rocket (120 sec).

Litchford, R. J.↗

NASA NTRS2001

Applications of the Method of Space-Time Conservation Element and the Solution Element to Unsteady Chemically Reactive Flows

This document reports the conclusion and findings of our research activities for this grant. The goal of the project is the development and application of the method of Space-Time Conservation Element and Solution Element, or the CE/SE method, to simulate chemically reacting flows. The product of this project will be a high-fidelity, time-accurate flow solver analyzing unsteady flow fields advanced propulsion concepts, including the low-emission turbojet engine combustion and flow fields of the Pulse Detonation Engines (PDE). Based on the documents and computer software of the CE/SE method that we have received from the CE/SE working group at NASA Lewis, we have focused our research effort on addressing outstanding technical issues related to the extension of the CE/SE method for unsteady, chemically reactive flows. In particular, we have made progresses in the following three aspects: (1) Derivation of the governing equations for reacting flows; (2) Numerical treatments of stiff source terms; and (3) Detailed simulations of ZND detonation waves.

Yu, Sheng-Tao↗

NASA NTRS2001

Performance Evaluation of the NASA GTX RBCC Flowpath

The NASA Glenn Research Center serves as NASAs lead center for aeropropulsion. Several programs are underway to explore revolutionary airbreathing propulsion systems in response to the challenge of reducing the cost of space transportation. Concepts being investigated include rocket-based combined cycle (RBCC), pulse detonation wave, and turbine-based combined cycle (TBCC) engines. The GTX concept is a vertical launched, horizontal landing, single stage to orbit (SSTO) vehicle utilizing RBCC engines. The propulsion pod has a nearly half-axisymmetric flowpath that incorporates a rocket and ram-scramjet. The engine system operates from lift-off up to above Mach 10, at which point the airbreathing engine flowpath is closed off, and the rocket alone powers the vehicle to orbit. The paper presents an overview of the research efforts supporting the development of this RBCC propulsion system. The experimental efforts of this program consist of a series of test rigs. Each rig is focused on development and optimization of the flowpath over a specific operating mode of the engine. These rigs collectively establish propulsion system performance over all modes of operation, therefore, covering the entire speed range. Computational Fluid Mechanics (CFD) analysis is an important element of the GTX propulsion system development and validation. These efforts guide experiments and flowpath design, provide insight into experimental data, and extend results to conditions and scales not achievable in ground test facilities. Some examples of important CFD results are presented.

Thomas, Scott R.↗

NASA NTRS2001

Laser Powered Launch Vehicle Performance Analyses

The purpose of this study is to establish the technical ground for modeling the physics of laser powered pulse detonation phenomenon. Laser powered propulsion systems involve complex fluid dynamics, thermodynamics and radiative transfer processes. Successful predictions of the performance of laser powered launch vehicle concepts depend on the sophisticate models that reflects the underlying flow physics including the laser ray tracing the focusing, inverse Bremsstrahlung (IB) effects, finite-rate air chemistry, thermal non-equilibrium, plasma radiation and detonation wave propagation, etc. The proposed work will extend the base-line numerical model to an efficient design analysis tool. The proposed model is suitable for 3-D analysis using parallel computing methods.

Chen, Yen-Sen↗

NASA NTRS1995

Preliminary Studies of a Pulsed Detonation Rocket Engine

In the new era of space exploration, there is a strong need for more efficient, cheaper and more reliable propulsion devices. With dramatic increase in specific impulse, the overall mass of fuel to be lifted into orbit is decreased, and this leads, in turn, to much lower mass requirements at lift-off, higher payload ratios and lower launch costs. The Pulsed Detonation engine (PDE) has received much attention lately due to its unique combination of simplicity, light-weight and efficiency. Current investigations focus principally on its use as a low speed, airbreathing engine, although other applications have also been proposed. Its use as a rocket propulsion device was first proposed in 1988 by the present authors. The superior efficiency of the Pulsed Detonation Rocket Engine (PDRE) is due to the near constant volume combustion process of a detonation wave. Our preliminary estimates suggest that the PDRE is theoretically capable of achieving specific impulses as high as 720 sec, a dramatic improvement over the current 480 sec of conventional rocket engines, making it competitive with nuclear thermal rockets. In addition to this remarkable efficiency, the PDRE may eliminate the need for high pressure cryogenic turbopumps, a principal source of failures. The heat transfer rates are also much lower, eliminating the need for nozzle cooling. Overall, the engine is more reliable and has a much lower weight. This paper will describe in detail the operation of the PDRE and calculate its performance, through numerical simulations. Engineering issues will be addressed and discussed, and the impact on mission profiles will also be presented. Finally, the performance of the PDRE using in-situ resources, such as CO and O2 from the martian atmosphere, will also be computed.

Cambier, Jean-Luc↗

NASA NTRS2003

Pulse Detonation Rocket Magnetohydrodynamic Power Experiment

The production of onboard electrical power by pulse detonation engines is problematic in that they generate no shaft power; however, pulse detonation driven magnetohydrodynamic (MHD) power generation represents one intriguing possibility for attaining self-sustained engine operation and generating large quantities of burst power for onboard electrical systems. To examine this possibility further, a simple heat-sink apparatus was developed for experimentally investigating pulse detonation driven MHD generator concepts. The hydrogen oxygen fired driver was a 90 cm long stainless steel tube having a 4.5 cm square internal cross section and a short Schelkin spiral near the head end to promote rapid formation of a detonation wave. The tube was intermittently filled to atmospheric pressure and seeded with a CsOH/methanol prior to ignition by electrical spark. The driver exhausted through an aluminum nozzle having an area contraction ratio of A*/A(sub zeta) = 1/10 and an area expansion ratio of A(sub zeta)/A* = 3.2 (as limited by available magnet bore size). The nozzle exhausted through a 24-electrode segmented Faraday channel (30.5 cm active length), which was inserted into a 0.6 T permanent magnet assembly. Initial experiments verified proper drive operation with and without the nozzle attachment, and head end pressure and time resolved thrust measurements were acquired. The exhaust jet from the nozzle was interrogated using a polychromatic microwave interferometer yielding an electron number density on the order of 10(exp 12)/cm at the generator entrance. In this case, MHD power generation experiments suffered from severe near-electrode voltage drops and low MHD interaction; i.e., low flow velocity, due to an inherent physical constraint on expansion with the available magnet. Increased scaling, improved seeding techniques, higher magnetic fields, and higher expansion ratios are expected to greatly improve performance.

Litchford, R. J.↗