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NASA NTRS1974

Combustion modeling for experimentation in a space environment

The merits of combustion experimentation in a space environment are assessed, and the impact of such experimentation on current theoretical models is considered. It is noted that combustion theory and experimentation for less than normal gravitational conditions are incomplete, inadequate, or nonexistent. Extensive and systematic experimentation in a space environment is viewed as essential for more adequate and complete theoretical models of such processes as premixed flame propagation and extinction limits, premixed flame propagation in droplet and particle clouds, ignition and autoignition in premixed combustible media, and gas jet combustion of unpremixed reactants. Current theories and models in these areas are described, and some combustion studies that can be undertaken in the Space Shuttle Program are proposed, including crossed molecular beam, turbulence, and upper pressure limit (of gases) studies.

Berlad, A. L.↗

NASA NTRS1998

LSPRAY: Lagrangian Spray Solver for Applications With Parallel Computing and Unstructured Gas-Phase Flow Solvers

Sprays occur in a wide variety of industrial and power applications and in the processing of materials. A liquid spray is a phase flow with a gas as the continuous phase and a liquid as the dispersed phase (in the form of droplets or ligaments). Interactions between the two phases, which are coupled through exchanges of mass, momentum, and energy, can occur in different ways at different times and locations involving various thermal, mass, and fluid dynamic factors. An understanding of the flow, combustion, and thermal properties of a rapidly vaporizing spray requires careful modeling of the rate-controlling processes associated with the spray's turbulent transport, mixing, chemical kinetics, evaporation, and spreading rates, as well as other phenomena. In an attempt to advance the state-of-the-art in multidimensional numerical methods, we at the NASA Lewis Research Center extended our previous work on sprays to unstructured grids and parallel computing. LSPRAY, which was developed by M.S. Raju of Nyma, Inc., is designed to be massively parallel and could easily be coupled with any existing gas-phase flow and/or Monte Carlo probability density function (PDF) solver. The LSPRAY solver accommodates the use of an unstructured mesh with mixed triangular, quadrilateral, and/or tetrahedral elements in the gas-phase solvers. It is used specifically for fuel sprays within gas turbine combustors, but it has many other uses. The spray model used in LSPRAY provided favorable results when applied to stratified-charge rotary combustion (Wankel) engines and several other confined and unconfined spray flames. The source code will be available with the National Combustion Code (NCC) as a complete package.

Raju, Manthena S.↗

NASA NTRS1980

Automated holographic drop-size analyzer

System analyzes drop-size distribution in liquid-droplet-spray combustion fields. Holographic camera takes "stop-motion" hologram of combustion volume; it is then viewed by vidicon camera connected to digital data-processing system that identifies particles or droplets, determining their size and count, and displays histogram of drop-size distribution in holographic field.

Feinstein, S. P.↗

NASA NTRS1983

Fundamental combustion experiments in microgravity

Technical bases for microgravity combustion experimentation are identified for important areas of study. The analyses which underlie the need for microgravity experiments of single and two-phase autoignitions, explosions, thermokinetic oscillations, and kinetic oscillations are emphasized. It is shown that turbulent combustion experiments at microgravity are needed and that heterogeneous kinetics are a centrally important process in a number of classic experiments.

Berlad, A. L.↗

NASA NTRS1987

Particle cloud kinetics in microgravity

Data related to particle-particle agglomeration/deaglomeration and particle-wall attachment are discussed. Particle-cluster cloud interactions and particle-cluster agglomeration/deagglomeration kinetics are studied. An apparatus designed and constructed for examining the agglomeration/deagglomeration effects for lycopodium under the conditions of an acoustically energized mixing process and of alpha-particle-induced deagglomerative processes is described; characteristic features and applications for the apparatus are examined. Requirements for combustion experimentation are discussed.

Berlad, A. L.↗

NASA NTRS1992

NLS propulsion - Government view

The paper discusses the technology development for the Space Transportation Main Engine (STME). The STME is a liquid oxygen/liquid hydrogen engine with 650,000 pounds of thrust, which may be flown in single-engine or multiple-engine configurations, depending upon the payload and mission requirements. The technological developments completed so far include a vacuum plasma spray process, the liquid interface diffusion bonding, and a thin membrane platelet technology for the combustion chamber fabrication; baseline designs for the hydrogen turbopump and the oxygen pump; and the engine control system. The family of spacecraft for which this engine is being developed includes a 20,000 pound payload to LEO and a 150,000 pound to LEO vehicle.

Smelser, Jerry W.↗

NASA NTRS1995

Advanced Small Rocket Chambers. Option 3: 110 1Bf Ir-Re Rocket, Volume 1

This report describes the AJ10-221, a high performance Iridium-coated Rhenium (Ir-Re) 110 lbf (490N) welded rocket chamber with 286:1 area ratio nozzle. This engine was designed, built, and hot fired for over 6 hours on this program. It demonstrated an I(s) of 321.8 sec, which is 10 sec higher than conventional 110 lbf silicide coated Cb chambers now in use. The approach used in this portion of the program was to demonstrate the performance improvement that can be made by the elimination of fuel film cooling and the use of a high temperature (4000F) (2200C) iridium-coated rhenium (Ir-Re) rocket chamber. Detailed thermal, performance, mechanical, and dynamic design analyses of the full engine were conducted by Aerojet. Two Ir-Re chambers were built to the Aerojet design by Ultramet, using the chemical vapor deposition (CVD) process. Incorporation of a secondary mixing device or Boundary Layer Trip (BLT) within the combustion chamber (Aerojet Patents 4882904 and 4936091) results in improvement in flow uniformity, and a significant life and performance increase. The 110 lbf engine design was verified in bolt-up hardware tests at sea level and altitude. The effects of injector design on performance were studied. Two duplicate injectors were fabricated matching the preferred design and were demonstrated to be interchangeable in operation. One of these units was fabricated matching the preferred design and was demonstrated to be interchangeable in operation. One of these units was welded into a flight type thruster which was tested for an accumulated duration of 22,590 sec in 93 firings, one of which was a continuous burn of two hours. A design deficiency in the C-103 nozzle near the Re-Cb transition joint was discovered, studied and corrected design has been prepared. Workhardening studies have been conducted to investigate methods for increasing the low yield strength of the Re in the annealed conditions. An advanced 490N high performance engine has been demonstrated which, when proven to be capable of withstanding launch vibration, is ready for flight qualification.

Jassowski, Donald M.↗

NASA NTRS1995

Advanced small rocket chambers. Option 3: 110 1bf Ir-Re rocket, volume 2

This is the second part of a two-part report that describes the AJ10-221, a high performance iridium-coated rhenium (Ir-Re) 110 lbf (490N) welded rocket chamber with 286:1 area ratio nozzle. This engine was designed, built, and hot fired for over 6 hours on this program. It demonstrated an I(s) of 321.8 sec, which is 10 sec higher than conventional 110 lbf silicide coated Cb chambers now in use. The approach used in this portion of the program was to demonstrate the performance improvement that can be made by the elimination of fuel film cooling and the use of a high temperature (4000 F) (2200 C) iridium-coated rhenium (Ir-Re) rocket chamber. Detailed thermal, performance, mechanical, and dynamic design analyses of the full engine were conducted by Aerojet. Two Ir-Re chambers were built to the Aerojet design by Ultramet, using the chemical vapor deposition (CVD) process. Incorporation of a secondary mixing device or Boundary Layer Trip (BLT) within the combustion chamber (Aerojet Patents 4882904 and 4936091) results in improvement in flow uniformity, and a significant life and performance increase. The 110 lbf engine design was verified in bolt-up hardware tests at sea level and altitude. The effects of injector design on performance were studied. Two duplicate injectors were fabricated matching the preferred design and were demonstrated to be interchangeable in operation. One of these units were welded into a flight type thruster which was tested for an accumulated duration of 22,590 sec in 93 firings, one of which was a continuous burn of two hours. A design deficiency in the C-103 nozzle near the Re-Cb transition joint was discovered, studied and corrected design has been prepared. Workhardening studies have been conducted to investigate methods for increasing the low yield strength of the Re in the annealed conditions. An advanced 490N high performance engine has been demonstrated which, when proven to be capable of withstanding launch vibration, is ready for flight qualification.

Jassowski, Donald M.↗

NASA NTRS1995

STS-64 Space Shuttle mission report

The STS-64 Space Shuttle Program Mission Report summarizes the Payload activities as well as the Orbiter, External Tank (ET), Solid Rocket Booster (SRB), Redesigned Solid Rocket Motor (RSRM), and the Space Shuttle main engine (SSME) systems performance during the sixty-fourth flight of the Space Shuttle Program and the nineteenth flight of the Orbiter vehicle Discovery (OV-103). In addition to the Orbiter, the flight vehicle consisted of an ET that was designated ET-66; three SSMEs that were designated as serial numbers 2031, 2109, and 2029 in positions 1, 2, and 3, respectively; and two SRB's that were designated Bl-068. The RSRM's that were installed in each SRB were designated as 360L041 A for the left SRB, and 360L041 B for the right SRB. The primary objective of this flight was to successfully perform the planned operations of the Lidar In-Space Technology Experiment (LITE), and to deploy the Shuttle Pointed Autonomous Research Tool for Astronomy (SPARTAN) -201 payload. The secondary objectives were to perform the planned activities of the Robot Operated Materials Processing System (ROMPS), the Shuttle Amateur Radio Experiment - 2 (SAREX-2), the Solid Surface Combustion Experiment (SSCE), the Biological Research in Canisters (BRIC) experiment, the Radiation Monitoring Equipment-3 (RME-3) payload, the Military Application of Ship Tracks (MAST) experiment, and the Air Force Maui Optical Site Calibration Test (AMOS) payload.

Fricke, Robert W., Jr.↗

NASA NTRS1998

Droplet Combustion Experiment (DCE)

The first space-based experiments were performed on the combustion of free, individual liquid fuel droplets in oxidizing atmospheres. The fuel was heptane, with initial droplet diameters ranging about from 1 mm to 4 mm. The atmospheres were mixtures of helium and oxygen, at pressures of 1.00, 0.50 and 0.25 bar, with oxygen mole fractions between 20% and 40%, as well as normal Spacelab cabin air. The temperatures of the atmospheres and of the initial liquid fuel were nominally 300 K. A total of 44 droplets were burned successfully on the two flights, 8 on the shortened STS-83 mission and 36 on STS-94. The results spanned the full range of heptane droplet combustion behavior, from radiative flame extinction at larger droplet diameters in the more dilute atmospheres to diffusive extinction in the less dilute atmospheres, with the droplet disappearing prior to flame extinction at the highest oxygen concentrations. Quasisteady histories of droplet diameters were observed along with unsteady histories of flame diameters. New and detailed information was obtained on burning rates, flame characteristics and soot behavior. The results have motivated new computational and theoretical investigations of droplet combustion, improving knowledge of the chemical kinetics, fluid mechanics and heat and mass transfer processes involved in burning liquid fuels.

Haggard, John B., Jr.↗

NASA NTRS1999

DRACO Flowpath Performance and Environments

The Advanced Space Transportation (AST) project office has challenged NASA to design, manufacture, ground-test and flight-test an axisymmetric, hydrocarbon-fueled, flight-weight, ejector-ramjet engine system testbed no later than 2005. To accomplish this, a multi-center NASA team has been assembled. The goal of this team, led by NASA-Marshall Space Flight Center (MSFC), is to develop propulsion technologies that demonstrate rocket and airbreathing combined-cycle operation (DRACO). Current technical activities include flowpath conceptual design, engine systems conceptual design, and feasibility studies investigating the integration and operation of the DRACO engine with a Lockheed D-21B drone. This paper focuses on the activities of the Flowpath Systems Product Development Team (PDT), led by NASA-Glenn Research Center (GRC) and supported by NASA-MSFC and TechLand Research, Inc. The objective of the Flowpath PDT at the start of the DRACO program was to establish a conceptual design of the flowpath aerodynamic lines, determine the preliminary performance, define the internal environments, and support the DRACO testbed concept feasibility studies. To accomplish these tasks, the PDT convened to establish a baseline flowpath concept. With the conceptual lines defined, cycle analysis tasks were planned and the flowpath performance and internal environments were defined. Additionally, sensitivity studies investigating the effects of inlet reference area, combustion performance, and combustor/nozzle materials selection were performed to support the Flowpath PDT design process. Results of these tasks are the emphasis of this paper and are intended to verify the feasibility of the DRACO flowpath and engine system as well as identify the primary technical challenges inherent in the flight-weight design of an advanced propulsion technology demonstration engine. Preliminary cycle performance decks were developed to support the testbed concept feasibility studies but are not discussed further in this paper.

Komar, D. R.↗

NASA NTRS2014

Pre-Test CFD for the Design and Execution of the Enhanced Injection and Mixing Project at NASA Langley Research Center

With the increasing costs of physics experiments and simultaneous increase in availability and maturity of computational tools it is not surprising that computational fluid dynamics (CFD) is playing an increasingly important role, not only in post-test investigations, but also in the early stages of experimental planning. This paper describes a CFD-based effort executed in close collaboration between computational fluid dynamicists and experimentalists to develop a virtual experiment during the early planning stages of the Enhanced Injection and Mixing project at NASA Langley Research Center. This projects aims to investigate supersonic combustion ramjet (scramjet) fuel injection and mixing physics, improve the understanding of underlying physical processes, and develop enhancement strategies and functional relationships relevant to flight Mach numbers greater than 8. The purpose of the virtual experiment was to provide flow field data to aid in the design of the experimental apparatus and the in-stream rake probes, to verify the nonintrusive measurements based on NO-PLIF, and to perform pre-test analysis of quantities obtainable from the experiment and CFD. The approach also allowed for the joint team to develop common data processing and analysis tools, and to test research ideas. The virtual experiment consisted of a series of Reynolds-averaged simulations (RAS). These simulations included the facility nozzle, the experimental apparatus with a baseline strut injector, and the test cabin. Pure helium and helium-air mixtures were used to determine the efficacy of different inert gases to model hydrogen injection. The results of the simulations were analyzed by computing mixing efficiency, total pressure recovery, and stream thrust potential. As the experimental effort progresses, the simulation results will be compared with the experimental data to calibrate the modeling constants present in the CFD and validate simulation fidelity. CFD will also be used to investigate different injector concepts, improve understanding of the flow structure and flow physics, and develop functional relationships. Both RAS and large eddy simulations (LES) are planned for post-test analysis of the experimental data.

Drozda, Tomasz G.↗

NASA NTRS2016

Development of Detonation Modeling Capabilities for Rocket Test Facilities: Hydrogen-Oxygen-Nitrogen Mixtures

The objective of the presented work was to develop validated computational fluid dynamics (CFD) based methodologies for predicting propellant detonations and their associated blast environments. Applications of interest were scenarios relevant to rocket propulsion test and launch facilities. All model development was conducted within the framework of the Loci/CHEM CFD tool due to its reliability and robustness in predicting high-speed combusting flow-fields associated with rocket engines and plumes. During the course of the project, verification and validation studies were completed for hydrogen-fueled detonation phenomena such as shock-induced combustion, confined detonation waves, vapor cloud explosions, and deflagration-to-detonation transition (DDT) processes. The DDT validation cases included predicting flame acceleration mechanisms associated with turbulent flame-jets and flow-obstacles. Excellent comparison between test data and model predictions were observed. The proposed CFD methodology was then successfully applied to model a detonation event that occurred during liquid oxygen/gaseous hydrogen rocket diffuser testing at NASA Stennis Space Center.

Allgood, Daniel C.↗

NASA NTRSDate not supplied

Benefits of Trash-to-Gas Versus Jettison of Waste Via Trash-Lock for Mars Transit

Human exploration missions to Mars pose difficulties due to the significant waste that will be generated during transit, which will need to be carried along or disposed of in some fashion. Waste removal from the spacecraft decreases the spacecraft’s mass as well as the associated logistic items necessary for storing the waste. A mission propellant analysis was performed to highlight the mass benefits that may be accessed via waste removal. The propellant mass savings were determined for different waste removal rates (2.9 – 11.6 kg/day) with the highest removal rate leading to the greatest propellant savings of 7,785 kg for an 850-day round-trip mission. Due to these benefits, two methods for waste reduction were studied for the 850-day Mars mission: Trash-to-Gas (TtG) and physical jettison via a trash-lock. The trash-to-gas methods considered were combustion, steam reforming, and pyrolysis, which convert waste into ventable gases (e.g., CO2, CO, CH4, etc.). Combustion and steam reforming require a co-reactant (O2 and/or H2O). Therefore, additional processing units or integration with the spacecraft’s environmental control and life support system (ECLSS) are required to facilitate recycle of the pertinent species. In contrast, pyrolysis is a purely thermal degradation process, which can operate as a standalone system; however, a lower percentage of waste is gasified with pyrolysis. The study herein compares standalone TtG (e.g., Advanced Organic Waste Gasifier, Plasma Pyrolysis, etc.), integrated TtG-ECLSS (e.g., Orbital Syngas Commodity Augmentation Reactor, Incineration/Gasification, etc.), and physical jettison. Each system’s mass, volume, power, and cooling requirements were compared via an equivalent system mass (ESM) analysis to ascertain potentially promising technologies that can achieve efficient waste removal while minimizing their own spacecraft load. This study highlights the advantages and disadvantages of the different waste management technologies and provides recommendations on the promising technologies based on the ESM metric and propellant mass savings.

Jettison↗

NASA NTRSDate not supplied

Metal Additive Manufacturing for Rocket Engines: Successes and Failures

NASA has been developing additive manufacturing (AM) for various technical and programmatic advantages for complex rocket engine and aerospace components. This maturation has focused on various AM processes, new alloys, characterizing material properties, developing standards, producing demonstrator parts, and integrating AM hardware in liquid rocket engines through test-fail-fix cycles, as well as application and dissemination of lessons learned of the AM lifecycle. The importance of proper AM processing was made evident in the failure of a Laser Powder Bed Fusion (L-PBF) copper-alloy combustion chamber during a hot-fire test due to a degraded material quality. The hot-fire test aimed to demonstrate high duty cycle under a risk-tolerant development project, where consequences of component failure would be minimal. However, the unintentional component failure emphasized the necessity of robust material characterization and rigorous process control procedures for the safe use of AM components in critical applications. This presentation provides an overview of the development failure, a discussion on the evaluation of the failed chamber and supplemental chambers produced at the same time, a representative material samples that included intentional build witness lines, and a summary of the key results and recommendations from the evaluations. NASA continues to approach AM processes and designs with a level of risk and acceptance of failures that is appropriate for the project objectives, with the overall goal of safe implementation of AM technology and transferring AM technology into commercial space applications. This presentation with provide critical awareness to the community lessons learned on proper implementation of AM.

L-PBF↗

NASA NTRSDate not supplied

Successes and Failures of Metal Additive Manufacturing for Rocket Engines

Additive Manufacturing↗

NASA NTRS1987

Liquid Engine Jet Atomization Workshop report

The workshop was oriented to disclose information and unsettled problems to understand the fundamental physical mechanism of the droplet formation process. Based on presentation and discussion of results, recommendations were made which should lead to associated future activities. To accomplish this task, existing observations and experiments, contributing to the basic knowledge, providing data for analytical concept verification, and forming a basis for empirical correlations were solicited. Advanced analytical modeling methods or results from specific studies were requested as well as the experience and advice from injector designers. All effort is directed to advance current analytical techniques, simulating the flow behavior downstream of the injection elements in a liquid rocket combustion chamber. Such a tool can be used to optimize injector designs with respect to short length weight savings, wall material protection, or large heat energy transport to a regenerative cooling fluid, while simultaneously achieving the maximum specific impulse in performance. The liquid atomization process also forms a sound basis for combustion instability analysis.

Gross, Klaus W.↗

NASA NTRS2004

[Analysis of Multiplatform CO (Carbon Monoxide) Measurements During Trace-P Mission]

Carbon monoxide is considered mission critical (TRACE-P NRA) because it is one of the gases involved in controlling the oxidizing power of the atmosphere and, as a tracer gas, is valuable in interpreting mission data sets. Carbon monoxide exhibits interannual differences, suggesting relatively short-term imbalances in sources and sinks. Sources of CO are dominated by fossil fuel combustion, biomass burning, and the photochemical oxidation of CH4 and nonmethane hydrocarbons while reaction with OH is believed to be the major sink for atmospheric CO, with additional losses due to soil uptake. Uncertainties in the magnitude and distribution of both sources and sinks remain fairly large however, and additional data are required to refine the global budget. Seasonal changes and a northern hemispheric latitudinal gradient have been described for a variety of Pacific basin sites through long-term monitoring of surface background levels. Latitudinal variations have also recently been described at upper tropospheric altitudes over a multi-year period by. TRACE-P will provide an aircraft survey of CO over the northern Pacific in the northern spring when CO concentrations are at their seasonal maximum in the northern hemisphere (NH) and at their seasonal minimum in the southern hemisphere (SH). Previous GTE missions, Le., PEM West-B and PEM Tropics-B, ground-based, and satellite observations (MAPS, April 1994) give us a general picture of the distribution of CO over the northern Pacific during this season. Based on these measurements, background CO levels over remote ocean areas are anticipated to be in the range of 110 - 180 ppbv, while those closer to the Asian continent may rise as high as 600 ppbv. These measurements also reveal high spatial variability (both horizontal and vertical) as well as temporal variations in CO over the area planned for the TRACE-P mission. This variability is a result of multiple CO sources, the meteorological complexity of transport processes, and the photochemical aging of air masses. The influence of biomass burning in the southern Pacific should be relatively small since the mission coincides with the southern tropical wet season when agricultural burning is at its seasonal low. The proposed CO measurements taken during TRACE-P should therefore largely be a function of the impact of various NH sources, primarily Asian and predominantly fossil fuel combustion and biomass burning. These processes are also major sources of many other atmospheric pollutants, consequently making accurate and precise CO measurements is one of the highest TRACE-P priorities [TRACE-P NRA]. The TRACE-P mission emphasizes the dual objectives of assessing the magnitude of the transport of chemically and radiatively important gases such as CO from Asia to the western Pacific, and determining how emissions change and are modified during this transport.

Pougatchev, Nikita S.↗