Results for “Enthalpy method” — Page 9

NASA NTRS2018

Calculation of Thermochemical Properties of Carbon-cluster Ablation Species

Carbon clusters and hydrocarbons are constituents of the pyrolysis gases injected into the boundary layer of a space vehicle with a carbonaceous heat shield. These molecules have absorption spectra in the VUV and UV region that match the emission spectra of atomic nitrogen and oxygen. Hence, they can potentially absorb the radiation impinging on the heat shield of the space vehicle. This paper studies the ground state thermochemical properties and low-lying excited electronic states of potential radiation absorbing molecules present in the boundary layer using ab initio quantum chemistry methods. These results provide a more accurate prediction of the radiative heat flux on the surface which can lead to improvement in the design of the thermal protection system.

enthalpy↗

NASA NTRS1977

Flight boundary layer transition measurements on a slender cone at Mach 20

Experimental flight boundary layer transition data have been obtained on a 3.962-m-long, 5 deg half-angle cone with an initial nose radius of 0.254 cm. The data were obtained during re-entry from altitudes of approximately 30.480 to 18.288 km at a free-stream Mach number of 20. The free-stream Reynolds number varied from 6.56 x 10 to the 6th/m to 52.5 x 10 to the 6th/m, and the total enthalpy from about 18.3 to 16.9 MJ/kg. The locations of the beginning and end of transition were determined by the intersection of curves faired throug the laminar, transitional, and turbulent heating-rate data. The temperature-history technique for determining transition as currently used (sharp break in curve) was shown to compare unfavorably with the heating-rate-distribution method. The heating-rate-history technique, which is proportional to the temperature derivative and consequently more sensitive to perturbations, gives better agreement with the heating-rate distribution transition results.

Wright, R. L.↗

NASA NTRS1994

LSENS: A General Chemical Kinetics and Sensitivity Analysis Code for homogeneous gas-phase reactions. Part 1: Theory and numerical solution procedures

LSENS, the Lewis General Chemical Kinetics and Sensitivity Analysis Code, has been developed for solving complex, homogeneous, gas-phase chemical kinetics problems and contains sensitivity analysis for a variety of problems, including nonisothermal situations. This report is part 1 of a series of three reference publications that describe LENS, provide a detailed guide to its usage, and present many example problems. Part 1 derives the governing equations and describes the numerical solution procedures for the types of problems that can be solved. The accuracy and efficiency of LSENS are examined by means of various test problems, and comparisons with other methods and codes are presented. LSENS is a flexible, convenient, accurate, and efficient solver for chemical reaction problems such as static system; steady, one-dimensional, inviscid flow; reaction behind incident shock wave, including boundary layer correction; and perfectly stirred (highly backmixed) reactor. In addition, the chemical equilibrium state can be computed for the following assigned states: temperature and pressure, enthalpy and pressure, temperature and volume, and internal energy and volume. For static problems the code computes the sensitivity coefficients of the dependent variables and their temporal derivatives with respect to the initial values of the dependent variables and/or the three rate coefficient parameters of the chemical reactions.

Radhakrishnan, Krishnan↗

NASA NTRS2000

Applications of Low Density Flow Techniques and Catalytic Recombination at the Johnson Space Center

The talk presents a brief background on defInitions of catalysis and effects associated with chemically nonequilibrium and low-density flows of aerospace interest. Applications of catalytic recombination on surfaces in dissociated flow are given, including aero heating on reentry spacecraft thermal protection surfaces and reflection of plume flow on pressure distributions associated with the space station. Examples include aero heating predictions for the X-38 test vehicle, the inlet of a proposed gas-sampling probe used in high enthalpy test facilities, and a parabolic body at angle of attack. The effect of accommodation coefficients on thruster induced pressure distributions is also included. Examples of tools used include simple aero heating formulas based on boundary layer solutions, an engineering approximation that uses axisymmetric viscous shock layer flow to simulate full three dimensional flow, full computational fluid dynamics, and direct simulation Monte-Carlo calculations. Methods of determining catalytic recombination rates in arc jet flow are discus ed. An area of catalysis not fully understood is the formation of single-wall carbon nanotubes (SWNT) with gas phase or nano-size metal particles. The Johnson Space Center is making SWNTs using both a laser ablation technique and an electric arc vaporization technique.

Scott, Carl D.↗

NASA NTRS1991

A drag measurement technique for free piston shock tunnels

A new technique is described for measuring drag with 100-microsecond rise time on a nonlifting model in a free piston shock tunnel. The technique involves interpretation of the stress waves propagating within the model and its support. A finite element representation and spectral methods are used to obtain a mean square optimal estimate of the time history of the aerodynamic loading. Thus, drag is measured instantaneously and the previous restriction caused by the mechanical time constant of balances is overcome. The effectiveness of the balance is demonstrated by measuring the drag on cones with 5 and 15 deg semi-vertex angles in nominally Mach 5.6 flow with stagnation enthalpies from 2.6 to 33 MJ/kg.

Sanderson, S. R.↗

NASA NTRS2019

Boundary Layer Stability and Laminar-Turbulent Transition Analysis with Thermochemical Nonequilibrium Applied to Martian Atmospheric Entry

As Martian atmospheric entry vehicles increase in size to accommodate larger payloads, transitional ow may need to be taken into account in the design of the heat shield in order to reduce heat shield mass. The mass of the Thermal Protection System (TPS) comprises a significant portion of the vehicle mass, and a reduction of this mass would result in fuel savings. The current techniques used to design entry shields generally assume fully turbulent flow when the vehicle is large enough to expect transitional flow, and while this worst-case scenario provides a greater factor of safety it may also result in overdesigned TPS and unnecessarily high vehicle mass. Greater accuracy in the prediction of transition would also reduce uncertainty in the thermal and aerodynamic loads. Stability analysis, using e(sup⁡ N) -based methods including Linear Stability Theory (LST) and the Parabolized Stability Equations (PSE), offers a physics-based method of transition prediction that has been thoroughly studied and applied in perfect gas flows, and to a more limited extent in reacting and nonequilibrium flows. These methods predict the amplification of a known disturbance frequency and allow identification of the most unstable frequency. Transition is predicted to occur at a critical amplification or N Factor, frequently determined through experiment and empirical correlations. The LAngley Stability and TRansition Analysis Code (LASTRAC), with modifications for thermochemically reacting flows and arbitrary gas mixtures, will be presented with LST results on a simulation of a high enthalpy CO2 gas wind tunnel test relevant to Martian atmospheric entry. The results indicate transition caused by modified Tollmien-Schlichting waves on the leeward side, which are predicted to be more stable and cause transition slightly downstream when thermochemical nonequilibrium is included in the stability analysis for the same mean flow solution.

Transition↗

NASA NTRS1990

Computation of three-dimensional turbulent boundary layers with heat transfer in a plane of symmetry using embedded wall-layer functions

In the calculation of turbulent boundary layers, a large number of mesh points are required to adequately resolve the intense variation in the velocity and enthalpy in the near-wall region. A substantial reduction in computational effort may be realized by representing the velocity and enthalpy profiles in the wall layer by analytical embedded functions. The effectively inviscid flow in the outer part of the boundary layer may then be resolved by employing a relatively coarser mesh. To obtain complete profiles, the outer numerical solution is matched asymptotically to the inner wall-layer analytical solution. To date, this approach has been restricted to two-dimensional flows; in the present study, a method which may be utilized for turbulent boundary layers with heat transfer in a plane of symmetry is developed as a first step in the application of the embedded-function method to full three-dimensional flows. The present method uses only about half as many mesh points as that required in a conventional procedure, which calculates the flow all the way to the wall, but there is no degradation in accuracy of the computed results.

Degani, A. T.↗

NASA NTRS1995

Measurements of scramjet thrust in shock tunnels

By using results obtained in tests on supersonic combustion of hydrogen in air, the conditions governing model size and operating pressure levels for shock tunnel experiments on models of flight vehicles with scramjet propulsion are established. It is seen that large models are required. The development of the stress wave force balance is then described, and its use as a method of measuring thrust/drag on such models is discussed. Test results on a simple, fully integrated scramjet model, with intakes, combustion chambers, thrust surfaces and exterior surfaces, using a 13 percent silane 87 percent hydrogen fuel mixture, showed that a steady state with thrust generation could be achieved within the shock tunnel test time, and the thrust could be measured. Results are presented for a range of stagnation enthalpies, and show that the scramjet model produces net positive thrust at velocities up to 2.4 km/sec.

Stalker, R. J.↗

NASA NTRS2000

Test Medium Effects on Ignition, Combustion and Flameholding Processes in Scramjet Combustors

The total temperatures (enthalpies) required to ground-test air-breathing (aero-propulsion) engines at high Mach number flight conditions can be achieved in a number of ways. Among these are: 1. Heat exchangers, including pre-heated ceramic beds. 2. direct electrical heating, e.g., arc discharge and resistance heaters. 3. Compression heating. 4. Shock heating, and 5. In-stream combustion, with oxygen replenishment to match air content. Each method has distinct advantages, disadvantages and limitations. All have a common characteristic of being designed for intermittent flow, due to the extreme energy required for continuous operation at simulated Mach numbers above about 3. All also distort the composition of atmospheric air to some degree, due to the high temperatures that occur in the plenum section prior to expansion of the flow to simulated flight conditions. In the case of in-stream combustion, the resulting test medium is commonly referred to as "vitiated air", being composed of oxygen, nitrogen and some fraction of combustion products.

Chinitz, W.↗

NASA NTRS2011

Ionic Liquids to Replace Hydrazine

A method for developing safe, easy-to-handle propellants has been developed based upon ionic liquids (ILs) or their eutectic mixtures. An IL is a binary combination of a typically organic cation and anion, which generally produces an ionic salt with a melting point below 100 deg C. Many ILs have melting points near, or even below, room temperature (room temperature ionic liquids, RTILs). More importantly, a number of ILs have a positive enthalpy of formation. This means the thermal energy released during decomposition reactions makes energetic ILs ideal for use as propellants. In this specific work, to date, a baseline set of energetic ILs has been identified, synthesized, and characterized. Many of the ILs in this set have excellent performance potential in their own right. In all, ten ILs were characterized for their enthalpy of formation, density, melting point, glass transition point (if applicable), and decomposition temperature. Enthalpy of formation was measured using a microcalorimeter designed specifically to test milligram amounts of energetic materials. Of the ten ILs characterized, five offer higher Isp performance than hydrazine, ranging between 10 and 113 seconds higher than the state-of-the-art propellant. To achieve this level of performance, the energetic cations 4- amino-l,2,4-triazolium and 3-amino-1,2,4-triazolium were paired with various anions in the nitrate, dicyanamide, chloride, and 3-nitro-l,2,4-triazole families. Protonation, alkylation, and butylation synthesis routes were used for creation of the different salts.

Koelfgen, Syri↗

NASA NTRS1987

The use of electrical discharge for ignition and control of combustion of solid propellants

As the first step of the study of the combustion control of solid propellants by electrical discharges, the effects of an arc discharge, which flows along the burning surface, on the burning rate and on the increase of enthalpy of the combustion product were investigated. For specially devised composite propellants, which are composed of Al and Teflon powders, it was shown that the combination can be controlled by an arc discharge; the combustion continues when the arc discharge is applied and is interrupted when the arc discharge breaks. In the present investigation, it was also shown that an arc discharge coupled with a high-frequency electrical discharge has potential as an effective ignition method for solid propellants. For the application of this type of combustion control to an ignitor for a solid propellant rocket motor or to a control rocket motor, this method lacks flexibility in the configuration scale and needs relatively high electric power at the present stage.

Tachibana, Takeshi↗

NASA NTRS1988

Electron heating and the potential jump across fast mode shocks

Two different methods were applied to determine the cross-shock potential jump in the de Hoffmann-Teller reference frame, using a data set that represented 66 crossings of the terrestrial bow shock and 14 interplanetary shocks observed by various ISEE spacecraft, and one crossing each of the Jovian bow shock and the Uranian bow shock made by the Voyager spacecraft. Results for estimates of the electrostatic potential based on an estimate of the jump in electron enthalpy correlated well with estimates based on Liouville's theorem, although the Liouville-determined values were systematically the higher of the two, suggesting that significant irreversible processes contribute to the shape of the downstream distribution. The potential jump corresponds to approximately 12-15 percent of the incident ion ram kinetic energy, and was found not to be controlled by the Mach number, plasma beta, shock geometry, or electron to ion temperature ratios.

Schwartz, Steven J.↗

NASA NTRSDate not supplied

Fluid System Stability Analysis Techniques

Fluid systems, or networks, consist of multiple components that work together to achieve some desired thermofluid state. For any generic application, this fluid state can be a combination of the fluid pressure, flow rate, enthalpy, or species concentration. Fluid system components, such as pumps and valves, are often governed by nonlinear differential equations, resulting in complex component-to-component interactions. System-level fluid network stability occurs when the flow through the system can maintain a steady-state solution in the presence of small perturbations, which depends on these component interactions. System instability, however, can go undetected until issues arise during integrated system testing. This presentation explores a method for system designers to think of the fluid network as an assembly of components, each with their own thermofluid surfaces of partial stability, called nullclines. The intersections of all nullclines yields system-level solutions, called equilibrium points. When designers define operating points, they are tuning system parameters so that these equilibrium points move to the desired location in the thermofluid state plane. However, linearization theory shows us that the dynamic behavior around these points can be unstable. The local stability of these equilibrium points can be assessed analytically with eigen-analysis, or numerically by propagating state-plane samples to construct a phase portrait. Investigating a phase portrait can help designers gain a qualitative understanding of a system’s dynamic performance. This understanding can then help inform requirement definitions, component selection, and operational procedures. This presentation includes an example of the phase portrait technique on a system featuring a centrifugal pump and a back-pressure regulator (BPR). Numerical modeling of this system suggests that equilibrium points on the left-hand side of the pump curve are dynamically unstable.

Fluid System↗

NASA NTRSDate not supplied

Fluid System Stability Analysis Techniques

Fluid systems, or networks, consist of multiple components that work together to achieve some desired thermofluid state. For any generic application, this fluid state can be a combination of the fluid pressure, flow rate, enthalpy, or species concentration. Fluid system components, such as pumps and valves, are often governed by nonlinear differential equations, resulting in complex component-to-component interactions. System-level fluid network stability occurs when the flow through the system can maintain a steady-state solution in the presence of small perturbations, which depends on these component interactions. System instability, however, can go undetected until issues arise during integrated system testing. This presentation explores a method for system designers to think of the fluid network as an assembly of components, each with their own thermofluid surfaces of partial stability, called nullclines. The intersections of all nullclines yields system-level solutions, called equilibrium points. When designers define operating points, they are tuning system parameters so that these equilibrium points move to the desired location in the thermofluid state plane. However, linearization theory shows us that the dynamic behavior around these points can be unstable. The local stability of these equilibrium points can be assessed analytically with eigen-analysis, or numerically by propagating state-plane samples to construct a phase portrait. Investigating a phase portrait can help designers gain a qualitative understanding of a system’s dynamic performance. This understanding can then help inform requirement definitions, component selection, and operational procedures. This presentation includes an example of the phase portrait technique on a system featuring a centrifugal pump and a back-pressure regulator (BPR). Numerical modeling of this system suggests that equilibrium points on the left-hand side of the pump curve are dynamically unstable.

Fluid System↗

NASA NTRS1991

Mach 4 and Mach 8 axisymmetric nozzles for a high-enthalpy shock tunnel

The performance of two axisymmetric nozzles which were designed to produce uniform, parallel flow with nominal Mach numbers of 4 and 8 is examined. A free-piston-driven shock tube was used to supply the nozzle with high-temperature, high-pressure test gas. The inviscid design procedure treated the nozzle expansion in two stages. Close to the nozzle throat, the nozzle wall was specified as conical and the gas flow was treated as a quasi-one-dimensional chemically-reacting flow. At the end of the conical expansion, the gas was assumed to be calorically perfect, and a contoured wall was designed (using method of characteristics) to convert the source flow into a uniform and parallel flow at the end of the nozzle. Performance was assessed by measuring Pitot pressures across the exit plane of the nozzles and, over the range of operating conditions examined, the nozzles produced satisfactory test flows. However, there were flow disturbances in the Mach 8 nozzle flow that persisted for significant times after flow initiation.

Jacobs, P. A.↗

NASA NTRS2005

An Approximate Ablative Thermal Protection System Sizing Tool for Entry System Design

A computer tool to perform entry vehicle ablative thermal protection systems sizing has been developed. Two options for calculating the thermal response are incorporated into the tool. One, an industry-standard, high-fidelity ablation and thermal response program was integrated into the tool, making use of simulated trajectory data to calculate its boundary conditions at the ablating surface. Second, an approximate method that uses heat of ablation data to estimate heat shield recession during entry has been coupled to a one-dimensional finite-difference calculation that calculates the in-depth thermal response. The in-depth solution accounts for material decomposition, but does not account for pyrolysis gas energy absorption through the material. Engineering correlations are used to estimate stagnation point convective and radiative heating as a function of time. The sizing tool calculates recovery enthalpy, wall enthalpy, surface pressure, and heat transfer coefficient. Verification of this tool is performed by comparison to past thermal protection system sizings for the Mars Pathfinder and Stardust entry systems and calculations are performed for an Apollo capsule entering the atmosphere at lunar and Mars return speeds.

Dec, John A.↗

NASA NTRS2006

An Approximate Ablative Thermal Protection System Sizing Tool for Entry System Design

A computer tool to perform entry vehicle ablative thermal protection systems sizing has been developed. Two options for calculating the thermal response are incorporated into the tool. One, an industry-standard, high-fidelity ablation and thermal response program was integrated into the tool, making use of simulated trajectory data to calculate its boundary conditions at the ablating surface. Second, an approximate method that uses heat of ablation data to estimate heat shield recession during entry has been coupled to a one-dimensional finite-difference calculation that calculates the in-depth thermal response. The in-depth solution accounts for material decomposition, but does not account for pyrolysis gas energy absorption through the material. Engineering correlations are used to estimate stagnation point convective and radiative heating as a function of time. The sizing tool calculates recovery enthalpy, wall enthalpy, surface pressure, and heat transfer coefficient. Verification of this tool is performed by comparison to past thermal protection system sizings for the Mars Pathfinder and Stardust entry systems and calculations are performed for an Apollo capsule entering the atmosphere at lunar and Mars return speeds.

Dec, John A.↗

NASA NTRS1986

Simulating AOTV heating environment in an arc jet

The application of arc-jet facilities to the simulation of the combined radiative and convective heating environments calculated for AOTV return missions is presented. The AOTV mission requirement and design configuration determine the magnitude of the radiative and convective heating rates. This paper considers methods for obtaining radiation from the high temperature plasma in arc-jet flows. A computer code that predicts arc-heater performance was used to determine the effect of gas mixtures on arc-heater performance, with the intent of increasing the gas temperature and resultant radiative emission. Numerical radiation solutions were compared with direct measurements of the radiative intensity from the gas cap. Convective heating measurements were made as well. The arc-jet was operated at AOTV enthalpy and pressure levels for these measurements to assess the simulation capabilities for AOTV material development/testing. A parametric study is presented that defines the capability of the Interaction Heating Facility arc jet at Ames to simulate both the radiative and the convective heating for current AOTV configurations.

Balboni, J.↗