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NASA NTRSDate not supplied

Effect of Nuclear Thermal Propulsion Engine Thrust and Isp Tradeoffs and Alternate Propellants on ΔV Budget and Architecture Mass for Crewed Mars Vehicles

This paper will provide an overview of a spacecraft architecture model developed at AMA as well as an insight into some of the results provided. Currently, preliminary analysis has been completed with this model on evaluating different thrust to weight ratios of Nuclear Thermal Propulsion based vehicles using hydrogen as the propellant by altering the number of engines in the engine block. Gravity losses are incorporated into this analysis as different thrust to weight ratios will provide different burn times. So far, this analysis has shown that in terms of the vehicle mass, it is best to trade the gravity losses for lower vehicle mass as the engines considered add significant dry mass to the vehicle. The final version of this paper will compare gravity losses and vehicle masses for vehicles using different engine masses as well as the ammonia propellant.

Master equipment list↗

NASA NTRS1985

Scramjet sidewall burning: Preliminary shock tunnel results

Experiments performed with a two dimensional model scramjet with particular emphasis on the effect of fuel injection from a wall are reported. Air low with a nominal Mach number of 3.5 and varied enthalpies was produced. It was found that neither hydrogen injection angle nor combustor divergence angle had any appreciable effect on thrust values while increased combustor length appeared to increase thrust levels. Specific impulse was observed to peak when hydrogen was injected at an equivalence ratio of about 2. Lowering the Mach number of the injected hydrogen at low equivalence ratios, less than 4, appeared to benefit specific impulse while hydrogen Mach number had little effect at higher equivalence ratios. When a 1:1 mixture by volume of nitrogen and oxygen is used instead of air as a test gas, it is found that hydrogen combustion is enhanced but only at high enthalpies.

Morgan, R. G.↗

NASA NTRS1993

Computational/experimental studies of isolated, single component droplet combustion

Isolated droplet combustion processes have been the subject of extensive experimental and theoretical investigations for nearly 40 years. The gross features of droplet burning are qualitatively embodied by simple theories and are relatively well understood. However, there remain significant aspects of droplet burning, particularly its dynamics, for which additional basic knowledge is needed for thorough interpretations and quantitative explanations of transient phenomena. Spherically-symmetric droplet combustion, which can only be approximated under conditions of both low Reynolds and Grashof numbers, represents the simplest geometrical configuration in which to study the coupled chemical/transport processes inherent within non-premixed flames. The research summarized here, concerns recent results on isolated, single component, droplet combustion under microgravity conditions, a program pursued jointly with F.A. Williams of the University of California, San Diego. The overall program involves developing and applying experimental methods to study the burning of isolated, single component droplets, in various atmospheres, primarily at atmospheric pressure and below, in both drop towers and aboard space-based platforms such as the Space Shuttle or Space Station. Both computational methods and asymptotic methods, the latter pursued mainly at UCSD, are used in developing the experimental test matrix, in analyzing results, and for extending theoretical understanding. Methanol, and the normal alkanes, n-heptane, and n-decane, have been selected as test fuels to study time-dependent droplet burning phenomena. The following sections summarizes the Princeton efforts on this program, describe work in progress, and briefly delineate future research directions.

Dryer, Frederick L.↗

NASA NTRS1972

A 10,000-gpm liquid hydrogen transfer system for the Saturn/Apollo program.

Brief description of the design and operation of the liquid hydrogen transfer system used to service the Saturn V launch vehicle. The cryogenic loading of the huge booster begins eight hours before the scheduled liftoff. The first three hours of fueling are spent in cold hydrogen gas conditioning of the fuel tank. The cold hydrogen gas is provided by vaporizing liquid hydrogen from the storage tank and routing the resultant gas through the fill system. Boil-off losses after loading are continuously replaced through control valves which are driven by a computer system. The liquid hydrogen transfer system is made up of a number of subsystems including the 850,000 gal storage tank whose boil-off losses amount to only 200 gal/day, the pressurization system, the burn pond for controlled disposal of hydrogen waste gas, the storage tank fill manifold, and the hazardous gas monitoring system. Some of the subsystems and components are redundant to provide a high degree of reliability.

Wybranowski, E., Jr.↗

NASA NTRS1990

On the use of external burning to reduce aerospace vehicle transonic drag

The external combustion of hydrogen to reduce the transonic drag of aerospace vehicles is currently being investigated. A preliminary analysis based on a constant pressure control volume is discussed. Results indicate that the specific impulse of the external burning process rivals that of a turbojet and depends on the severity of the initial base drag as well as on the Mach flight number and the equivalence ratio. A test program was conducted to investigate hydrogen-air flame stability at the conditions of interest and to demonstrate drag reduction on a simple expansion ramp. Initial test results are presented and compared with the control-volume analysis. The expansion ramp surface pressure coefficient showed little variation with fuel pressure and altitude, in disagreement with the analysis. Flame stability results were encouraging and indicate that stable combustion is possible over an adequate range of conditions. Facility interference and chemical kinetics phenomena that make interpretation of subscale ground test data difficult are discussed.

Trefny, Charles J.↗

NASA NTRS1990

On the use of external burning to reduce aerospace vehicle transonic drag

Trefny, Charles J.↗

NASA NTRS1991

Combustion of unsupported methanol/dodecanol mixture droplets at low gravity

Experiments are reported on the combustion of free or unsupported methanol/dodecanol mixture droplets in a reduced gravity environment. The experiments were conducted in a drop tower. The Grashof number (based on drop diameter) in the moving frame of reference was on the order of 10 exp -5 for the period of droplet burning. Combustion of the mixture droplets was characterized by a multi-stage process indicative of diffusion controlled burning. The flame was not visible during the first stage because of the lighting used, but was visible (and spherical) in the final stage. The results suggested that the mixture droplets did not extinguish. By contrast, methanol droplets consistently extinguished. Microexplosions were not observed for mixtures with initial concentrations of equal volumes of methanol and dodecanol. However, most of the droplets containing initially 75 percent (by volume) of methanol exploded. The initiation of microexplosion, as defined by the occurrence of homogeneous nucleation within the droplet, is explained by comparing the calculated nonlinear variation of superheat limit with composition for a methanol/dodecanol mixture with the boiling point of dodecanol.

Yang, J. C.↗

NASA NTRS2002

Measurement and Computation of Supersonic Flow in a Lobed Diffuser-Mixer for Trapped Vortex Combustors

The trapped vortex combustor (TVC) pioneered by Air Force Research Laboratories (AFRL) is under consideration as an alternative to conventional gas turbine combustors. The TVC has demonstrated excellent operational characteristics such as high combustion efficiency, low NO(x) emissions, effective flame stabilization, excellent high-altitude relight capability, and operation in the lean-burn or rich burn-quick quench-lean burn (RQL) modes of combustion. It also has excellent potential for lowering the engine combustor weight. This performance at low to moderate combustor mach numbers has stimulated interest in its ability to operate at higher combustion mach number, and for aerospace, this implies potentially higher flight mach numbers. To this end, a lobed diffuser-mixer that enhances the fuel-air mixing in the TVC combustor core was designed and evaluated, with special attention paid to the potential shock system entering the combustor core. For the present investigation, the lobed diffuser-mixer combustor rig is in a full annular configuration featuring sixfold symmetry among the lobes, symmetry within each lobe, and plain parallel, symmetric incident flow. During hardware cold-flow testing, significant discrepancies were found between computed and measured values for the pitot-probe-averaged static pressure profiles at the lobe exit plane. Computational fluid dynamics (CFD) simulations were initiated to determine whether the static pressure probe was causing high local flow-field disturbances in the supersonic flow exiting the diffuser-mixer and whether shock wave impingement on the pitot probe tip, pressure ports, or surface was the cause of the discrepancies. Simulations were performed with and without the pitot probe present in the modeling. A comparison of static pressure profiles without the probe showed that static pressure was off by nearly a factor of 2 over much of the radial profile, even when taking into account potential axial displacement of the probe by up to 0.25 in. (0.64 cm). Including the pitot probe in the CFD modeling and data interpretation lead to good agreement between measurement and prediction. Graphical inspection of the results showed that the shock waves impinging on the probe surface were highly nonuniform, with static pressure varying circumferentially among the pressure ports by over 10 percent in some cases. As part of the measurement methodology, such measurements should be routinely supplemented with CFD analyses that include the pitot probe as part of the flow-path geometry.

Brankovic, Andreja↗

NASA NTRS1985

Preliminary analysis of selected gas dynamic problems

The VAST computer code was used to analyze SSME main combustion chamber start-up transients and the IUS flow field for a damaged nozzle was investigated to better understand the gas dynamic considerations involved in vehicle problems, the effect of start transients on the nozzle flow field for the SSME, and the possibility that a damaged nozzle could account for the acceleration anomaly noted on IUS burn. The results obtained were compared with a method of characteristics prediction. Pressure solutions from both codes were in very good agreement and the Mach number solution on the nozzle centerline deviates substantially for the high expansions for the SSME. Since this deviation was unexpected, the phenomenon is being further examined.

Prozan, R. J.↗

NASA NTRS2015

Affordable Development and Demonstration of a Small NTR Engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

The Nuclear Thermal Rocket (NTR) represents the next evolutionary step in cryogenic liquid rocket engines. Deriving its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core, the NTR can generate high thrust at a specific impulse of approx. 900 seconds or more - twice that of today's best chemical rockets. In FY'11, as part of the AISP project, NASA proposed a Nuclear Thermal Propulsion (NTP) effort that envisioned two key activities - "Foundational Technology Development" followed by system-level "Technology Demonstrations". Five near-term NTP activities identified for Foundational Technology Development became the basis for the NCPS project started in FY'12 and funded by NASA's AES program. During Phase 1 (FY'12-14), the NCPS project was focused on (1) Recapturing fuel processing techniques and fabricating partial length "heritage" fuel elements for the two candidate fuel forms identified by NASA and the DOE - NERVA graphite "composite" and the uranium dioxide (UO2) in tungsten "cermet". The Phase 1 effort also included: (2) Engine Conceptual Design; (3) Mission Analysis and Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable and Sustainable NTP Development Strategy. During FY'14, a preliminary plan for DDT&E was outlined by GRC, the DOE and industry for NASA HQ that involved significant system-level demonstration projects that included GTD tests at the NNSS, followed by a FTD mission. To reduce development costs, the GTD and FTD tests use a small, low thrust (approx. 7.5 or 16.5 klbf) engine. Both engines use graphite composite fuel and a "common" fuel element design that is scalable to higher thrust (approx. 25 klbf) engines by increasing the number of elements in a larger diameter core that can produce greater thermal power output. To keep the FTD mission cost down, a simple "1-burn" lunar flyby mission was considered along with maximizing the use of existing and flight proven liquid rocket and stage hardware (e.g., from the RL10-B2 engine and Delta Cryogenic Second Stage) to further ensure affordability. This paper provides a preliminary NASA, DOE and industry assessment of what is required - the key DDT&E activities, development options, and the associated schedule - to affordably build, ground test and fly a small NTR engine and stage within a 10-year timeframe.

Nuclear rocket engines↗

NASA NTRSDate not supplied

Copernicus-LinCov (COPCOV) Software Integration in Support of Robust Trajectory Optimization

Robust trajectory optimization is the process of optimizing a trajectory while accounting for system uncertainty due to a variety of potential error sources. This work highlights the development and features of a novel tool known as CopCov to support robust trajectory optimization efforts. CopCov acts as an interface between Copernicus, a generalized trajectory design and optimization tool, and LinCov, a linear covariance analysis tool. By having a direct interface between these two software packages, Copernicus can receive covariance information from LinCov through a direct feedback loop, thus enabling optimization of a trajectory that is robust to trajectory dispersions and navigation errors. This paper details the architecture of CopCov and its flexibility to operate under varying configurations, including with both tools running locally or alternatively with the tools communicating via a remote connection. Additionally, the CopCov tool is demonstrated on a simple Hohmann transfer reference trajectory with varying numbers of Trajectory Correction Maneuvers (TCMs) and varying problem formulations. This example scenario is used to highlight how the inclusion of the CopCov interface affects burn placement of both major burns and minor burns (i.e., TCMs) in the optimized solution. Results are compared against analytical solutions and against a Genetic Algorithm (GA) optimizer for independent verification and validation.

Copernicus↗

NASA NTRS1995

Hot bottom burning in asymptotic giant branch stars and its effect on oxygen isotopic abundances

A self-consistent calculation of asymptotic giant branch (AGB) evolution was carried out, including nucleosynthesis at the base of the convective envelope (hot bottom burning). Hot bottom burning was found to occur for stars between approximately 4.5 and approximately 7 solar mass, producing envelopes with O-18/O-16 less than or equal to 10(exp -6) and 10(exp -3) approximately less than or equal O-17/O-16 approximately less than or equal to 10(exp -1). The O-17 abundance depends sensitively on the nuclear O-17-destruction rate; this rate is only loosely constrained by the requirement that first and second dredge-up models match O-isotope observations of red giant branch (RGB) stars (Boothroyd, Sackmann, & Wasserburg 1994). In some cases, high mass-loss rates can terminate hot bottom burning before further O-17 enrichment takes place or even before all O-18 is destroyed. These predictions are in accord with the very limited stellar observations of J type carbon stars on the AGB and with some of the circumstellar Al2O3 grains from meteorites. In contrast, precise data from a number of grains and data from most low-mass S and C AGB stars (approximately less than 1.7 solar mass) lie in a region of the O-18/O-16 versus O-17/O-16 diagram that is not accessible by first and second dredge-up or by hot bottom burning. We conclude that for AGB stars, the standard models of stellar evolution are not in accord with these observations. We surmise that an additional mixing mechanism must exist that transports material from the cool bottom of the stellar convective envelope to a depth at which O-18 is destroyed. This 'cool bottom processing' mechanism on the AGB is similar to extra mixing mechanisms proposed to explain the excess C-13 (and depleted C-12) observed in the earlier RGB stage of evolution and the large Li-7 depletion observed in low-mass main-sequence stars.

Boothroyd, Arnold I.↗

NASA NTRS1993

Numerical models of jet disruption in cluster cooling flows

We present a coherent picture for the formation of the observed diverse radio morphological structures in dominant cluster galaxies based on the jet Mach number. Realistic, supersonic, steady-state cooling flow atmospheres are evolved numerically and then used as the ambient medium through which jets of various properties are propagated. Low Mach number jets effectively stagnate due to the ram pressure of the cooling flow atmosphere while medium Mach number jets become unstable and disrupt in the cooling flow to form amorphous structures. High Mach number jets manage to avoid disruption and are able to propagate through the cooling flow.

Loken, Chris↗

NASA NTRS1994

Analysis of tropospheric aerosol number density for aerosols of 0.2- to 3-micrometers diameter: Central and northeastern Canada

NASA's Atmospheric Boundary Layer Experiment conducted during the summer of 1990 focused on the distribution of trace species in central and northeastern Canada (altitudes less than 6 km) and the importance of surface sources/sinks, local emissions, distant transport, tropospheric/stratospheric exchange. Aircraft flights were based from North Bay, Ontario, and Goose Bay, Labrador, Canada. As part of the aircraft measurements, aerosol number density (0.2- to 3-micrometers diameter) was measured using an optical laser technique. Results show that summertime aerosol budgets of central and northeastern Canada can be significantly impacted by the transport of pollutants from distant source regions. Biomass burning in Alaska and western and central Canada exerts major influences on regional aerosol budgets. Urban emissions transported from the U.S./Canadian border regions are also important. Aerosol enhancements (mixed layer and free troposphere) were most prevalent in air with carbon monoxide mixing ratios greater than 110 parts per billion by volume (ppbv). When data were grouped as to the source of the air (5-day back trajectories) either north or south of the polar jet, aerosol number density in the mixed layer showed a tendency to be enhanced for air south of the jet relative to north of the jet. However, this difference was not observed for measurements at the higher altitudes (4 to 6 km). For some flights, mixed layer aerosol number densities were greater than 100 higher than free-tropospheric values (3- to 6-km altitude). The majority of the observed mixed layer enhancement was associated with transport of effluent-rich air into the Canadian regions. Aerosol emissions from natural Canadian ecosystems were relatively small when compared to transport.

Gregory, Gerald L.↗

NASA NTRS1984

New technique for calibrating hydrocarbon gas flowmeters

A technique for measuring calibration correction factors for hydrocarbon mass flowmeters is described. It is based on the Nernst theorem for matching the partial pressure of oxygen in the combustion products of the test hydrocarbon, burned in oxygen-enriched air, with that in normal air. It is applied to a widely used type of commercial thermal mass flowmeter for a number of hydrocarbons. The calibration correction factors measured using this technique are in good agreement with the values obtained by other independent procedures. The technique is successfully applied to the measurement of differences as low as one percent of the effective hydrocarbon content of the natural gas test samples.

Singh, J. J.↗

NASA NTRS1986

Optical and probe determination of soot concentrations in a model gas turbine combustor

An experimental program was conducted to track the variation in soot loading in a generic gas turbine combustor. The burner is a 12.7-cm dia cylindrical device consisting of six sheet-metal louvers. Determination of soot loading along the burner length is achieved by measurement at the exit of the combustor and then at upstream stations by sequential removal of liner louvers to shorten burner length. Alteration of the flow field approaching and within the shortened burners is minimized by bypassing flow in order to maintain a constant linear pressure drop. The burner exhaust flow is sampled at the burner centerline to determine soot mass concentration and smoke number. Characteristic particle size and number density, transmissivity of the exhaust flow, and local radiation from luminous soot particles in the exhaust are determined by optical techniques. Four test fuels are burned at three fuel-air ratios to determine fuel chemical property and flow temperature influences. Particulate concentration data indicate a strong oxidation mechanism in the combustor secondary zone, though the oxidation is significantly affected by flow temperature. Soot production is directly related to fuel smoke point.

Eckerle, W. A.↗

NASA NTRS1932

Flame Movement and Pressure Development in an Engine Cylinder

This investigation describes a visual method for making stroboscopic observations, through a large number of small windows, of the spread of flame throughout the combustion chamber of a gasoline engine. Data, secured by this method on a small engine burning gaseous fuels, are given to show the effects of mixture ratio, spark advance, engine speed, charge density, degree of dilution, compression ratio, and fuel composition on flame movement in the cylinder. Partial indicator diagrams showing pressure development during the combustion period are included. Although present knowledge is not sufficient to permit qualitative evaluation of the separate effects on flame movement of chemical reaction velocity, thermal expansion of burned gases, resonance, turbulence, and piston movement, the qualitative influence of certain of these factors on some of the diagrams is indicated.

Marvin, Charles F , Jr↗

NASA NTRS1995

Combustion of interacting droplet arrays in a microgravity environment

This research program involves the study of one and two dimensional arrays of droplets in a buoyant-free environment. The purpose of the work is to extend the database and theories that exist for single droplets into the regime where droplet interactions are important. The eventual goal being to use the results of this work as inputs to models on spray combustion where droplets seldom burn individually; instead the combustion history of a droplet is strongly influenced by the presence of the neighboring droplets. Throughout the course of the work, a number of related aspects of isolated droplet combustion have also been investigated. This paper will review our progress in microgravity droplet array combustion, advanced diagnostics (specifically L2) applied to isolated droplet combustion, and radiative extinction large droplet flames. A small-scale droplet combustion experiment being developed for the Space Shuttle will also be described.

Dietrich, Daniel L.↗