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At least 343 records · Page 19

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NASA NTRS2023

Variably Premixed Rotating Detonation Engine for Evaluation of Detonation Cycle Dynamics

A variably premixed rotating detonation engine using gaseous hydrogen and air reactants is introduced to enable investigation of key cycle processes while varying the homogeneity of the reactant inlet conditions. Two chamber configurations are investigated, the first with reactants filling the entire span of a straight annular channel and the second with a slightly larger channel and a backward-facing step. The first configuration permits both premixed and non-premixed fuel injection, enabling mixing quality modulation. The second configuration is operated only at fully premixed conditions. Operating modes and detonation wave speeds are characterized using exhaust-plume imaging, while the chamber heat release field is captured by transverse imaging through a transparent outer body. Tests using the first configuration were characterized by unstable counterpropagating modes with low detonation wave speeds regardless of the state of premixing, while the second configuration rendered single-wave behavior with wave speeds up to 86% of the Chapman–Jouguet velocity. Comparisons with a simple computational fluid dynamics model of the second configuration indicate that reactant preheating significantly influences the detonation wave topology, highlighting the potential utility of the test platform for isolating key physics associated with the effects of reactant premixing, preheating, and chamber geometry on rotating detonation engine operation.

Detonation↗

NASA NTRS1999

A Theory of Oscillating Edge Flames

It has been known for some years that when a near-limit flame spreads over a liquid pool of fuel, the edge of the flame can oscillate relative to a frame moving with the mean speed. Each period of oscillation is characterized by long intervals of modest motion during which the edge gases radiate like those of a diffusion flame, punctuated by bursts of rapid advance during which the edge gases radiate like those in a deflagration. Substantial resources have been brought to bear on this issue within the microgravity program, both experimental and numerical. It is also known that when a near-asphyxiated candle-flame burns at zero gravity, the edge of the (hemispherical) flame can oscillate violently prior to extinction. Thus a web-surfer, turning to the NASA web-site at http://microgravity.msfc.nasa.gov, and following the trail combustion science/experiments/experimental results/candle flame, will find photographs and a description of candle burning experiments carried out on board both the Space-shuttle and the Russian space station Mir. A brief report can also be found in the proceedings of the Fourth Workshop. And recently, in a third microgravity program, the leading edge of the flame supported by injection of ethane through the porous surface of a plate over which air is blown has been found to oscillate when conditions are close to blow-off. A number of important points can be made with respect to these observations: It is the edge itself which oscillates, advancing and retreating, not the diffusion flame that trails behind the edge; oscillations only occur under near limit conditions; in each case the Lewis number of the fuel is significantly larger than 1; and because of the edge curvature, the heat losses from the reacting edge structure are larger than those from the trailing diffusion flame. We propose a general theory for these oscillations, invoking Occam's 'Law of Parsimony' in an expanded form, to wit: The same mechanism is responsible for the oscillations in all three experiments; and no new mechanism is invoked (Occam's original 'Razor'). Such a strategy eliminates Marangoni effects as the source, for these are absent in the second and third experiments. And it eliminates arguments that point to numerically predicted gas eddies as the source, a new mechanism, unelucidated. Indeed, we hypothesize that the essential driving mechanism for the instability is a combination of large Lewis number and heat losses from the reacting structure near the flame edge. Instabilities driven by these mechanisms are commonplace in 1D configurations. Chemical reactor theory, for example, leads to system responses which mimic the response of the candle flame - steady flame, oscillations, extinction. In a combustion context, oscillating instabilities were first reported for diffusion flames in a theoretical study by Kirkby and Schmitz, and here also the instabilities are associated with near-extinction conditions, large Lewis numbers, and heat losses. And deflagrations will oscillate if the Lewis number is large enough, oscillations that are exacerbated when heat losses are present, whether global or to a surface.

Buckmaster, J.↗

NASA NTRS1977

Fracture control of H-O engine components

An investigation was made to obtain the material characterization and fatigue crack propagation data necessary to establish the salient characteristics of a Ti-6Al-2.5Sn(ELI) alloy fuel pump impeller to be used in a cryogenic service environment. Testing variables considered were: coupon orientation, frequency, load range ratio, and temperature. Data analysis correlated crack propagation data from conventional laboratory coupons with data from a parallel sided rotating disk used to model rotor stresses. Four major design recommendations when bore regions of fuel pump impellers to be operated in cryogenic environments are to be relatively highly stressed are discussed.

Ryder, J. T.↗

NASA NTRS1989

Physical and chemical characteristics of cenospheres from the combustion of heavy fuel oil

Photomicrography of particle cross sections, measurements of density, porosity, and surface area, and determinations of chemical compositions, have been used in conjunction with SEM of surface structure to characterize cenospheres generated by combustion of residual oil in a steam power plant. Large and small cenospheres, which respectively fall into the 100-200 and small 20-40 micron range, are spheroidal and hollow, with at least one blowhole; outer/inner diameter ratios for the shells are of the order of 1.3-1.4. Typically, a cenosphere contains only about 18 vol pct solid material. The presence of S, Fe, Na, and V in substantial concentrations presage high temperature heat exchanger surface corrosion problems due to cenosphere deposition.

Clayton, R. M.↗

NASA NTRS1991

Design of multihundredwatt DIPS for robotic space missions

Design of a dynamic isotope power system (DIPS) general purpose heat source (GPHS) and small free piston Stirling engine (FPSE) is being pursued as a potential lower cost alternative to radioisotope thermoelectric generators (RTG's). The design is targeted at the power needs of future unmanned deep space and planetary surface exploration missions ranging from scientific probes to SEI precursor missions. These are multihundredwatt missions. The incentive for any dynamic system is that it can save fuel which reduces cost and radiological hazard. However, unlike a conventional DIPS based on turbomachinery converions, the small Stirling DIPS can be advantageously scaled to multihundred watt unit size while preserving size and weight competitiveness with RTG's. Stirling conversion extends the range where dynamic systems are competitive to hundreds of watts (a power range not previously considered for dynamic systems). The challenge of course is to demonstrate reliability similar to RTG experience. Since the competative potential of FPSE as an isotope converter was first identified, work has focused on the feasibility of directly integrating GPHS with the Stirling heater head. Extensive thermal modeling of various radiatively coupled heat source/heater head geometries were performed using data furnished by the developers of FPSE and GPHS. The analysis indicates that, for the 1050 K heater head configurations considered, GPHS fuel clad temperatures remain within safe operating limits under all conditions including shutdown of one engine. Based on these results, preliminary characterizations of multihundred watt units were established.

Bents, D. J.↗

NASA NTRS2000

Defect Characterization in a Thin Walled Composite RP-1 Tank: A Case Study

A full scale thin walled composite tank, designed and fabricated for the storage of pressurized RP- I rocket fuel, was fully inspected with digital infrared thermography (IR) during assembly and prior to proof testing. The tank featured a "pill capsule" design with the equatorial bondline being overwrapped on both the inner and outer surfaces. A composite skirt was bonded to the aft dome of the tank to serve as a structural support when the tank was stood on end in service. Numerous anomalies were detected and mapped prior to proof testing, some along bondlines and some scattered throughout the acreage. After the tank was intentionally burst, coupons were cut from the regions including thermographic anomalies. These coupons were again inspected thermographically to document the growth of any indications due to proof testing. Ultrasonic inspections (UT) were also performed on the coupons for comparison to thermography. Several coupons were dissected and micrographed. Relationships between IR and UT indications and the physical nature of the dissected material are presented.

Langsing, Matthew D.↗

NASA NTRS2002

Characterization of Subsystems for a WB-003 Single Stage Shuttle

Subsystems for an all oxygen-hydrogen-single-stage shuttle are characterized for a vehicle designated WB-003. Features of the vehicle include all-electric actuation, fiber optics for information circuitry, fuel cells for power generation, and extensive use of composites for structure. The vehicle is sized for the delivery of a 25,000 lb. payload to a space station orbit without crew. When crew are being delivered, they are carried in a module in the payload bay with escape and manual override capabilities. The underlying reason for undertaking this task is to provide a framework for the study of the operations costs of the newer shuttles.

MacConochie, Ian O.↗

NASA NTRS2011

Hyperthermal Environments Simulator for Nuclear Rocket Engine Development

An arc-heater driven hyperthermal convective environments simulator was recently developed and commissioned for long duration hot hydrogen exposure of nuclear thermal rocket materials. This newly established non-nuclear testing capability uses a high-power, multi-gas, wall-stabilized constricted arc-heater to produce hightemperature pressurized hydrogen flows representative of nuclear reactor core environments, excepting radiation effects, and is intended to serve as a low-cost facility for supporting non-nuclear developmental testing of hightemperature fissile fuels and structural materials. The resulting reactor environments simulator represents a valuable addition to the available inventory of non-nuclear test facilities and is uniquely capable of investigating and characterizing candidate fuel/structural materials, improving associated processing/fabrication techniques, and simulating reactor thermal hydraulics. This paper summarizes facility design and engineering development efforts and reports baseline operational characteristics as determined from a series of performance mapping and long duration capability demonstration tests. Potential follow-on developmental strategies are also suggested in view of the technical and policy challenges ahead. Keywords: Nuclear Rocket Engine, Reactor Environments, Non-Nuclear Testing, Fissile Fuel Development.

Litchford, Ron J.↗

NASA NTRS2018

Swirl-Venturi Lean Direct Injection Combustion Technology for Low-NOx Aero Gas Turbine Engines

This paper summarizes research on the lean direct injection (LDI) combustor concept for aero-gas turbine combustors. The focus of this paper is one particular family of lean direct injection designs, swirl-venturi lean direct injection (SV-LDI). SV-LDI is characterized by the airpath: an air swirler followed by a converging-diverging venturi. For most SV-LDI configurations, a fuel injector is inserted through the center of the air swirler, with the fuel injector tip at or near the venturi throat. Several design variables were studied. These included fuel injector tip location, air swirler blade thickness, air swirler blade angle, and fuel-air mixer size. Moving the fuel injector tip slightly upstream or downstream of the venturi throat has at most a small impact on NOx emissions. Changing the blade thickness also does not affect NOx emissions. Changing the swirler blade angle has a significant effect on NOx emissions. Decreasing swirler blade angle, and thus decreasing swirl number, decreases the NOx emissions at lower flame temperatures (below about 1800 K). However, the slope of the NOx vs. flame temperature curve is higher for lower swirl numbers. Finally, decreasing the fuel-air mixer size initially decreases NOx emissions. However, there may be an optimum fuel-air mixer size below which NOx emissions do not continue to decrease.

lean direct injection↗

NASA NTRS2019

Materials and Processes for New Propulsion Systems with Reduced Environmental Impact

Aeronautics research at NASA Glenn Research Center includes development, characterization and modeling of high temperature, lightweight materials and fabrication processes for aircraft propulsion systems with increased efficiency and reduced emissions, fuel burn and noise. Current propulsion materials research includes Ceramic Matrix Composites and Environmental Barrier Coatings, Polymer Matrix Composites and Additive Manufacturing processes. This presentation will summarize recent progress and plans in these areas.

Propulsion↗

NASA NTRSDate not supplied

Assessing Urban Heat in the Cincinnati and Covington Area to Inform Local Decision-Makers and Environmental Justice Initiatives

The Urban Heat Island effect is a phenomenon characterized by urban areas experiencing temperatures that are, on average, warmer than surrounding suburban and rural regions. Urban Heat Islands are fueled by expansive impervious surfaces, vehicle emissions, and insufficient urban green space. Densely populated urban centers like the Cincinnati, Ohio and Covington, Kentucky area can experience negative health impacts due to the Urban Heat Island. NASA DEVELOP partnered with Groundwork USA and Groundwork Ohio River Valley to combine environmental education and outreach with technical capacity building in NASA Earth observations. The DEVELOP team used Landsat 5 TM and ISS ECOSTRESS to calculate daytime and nighttime land surface temperature anomalies. The team found that the Cincinnati and Covington area was 8.32°F warmer during the day and 4.97°F warmer at night compared to non-urban areas. The team also ran the Natural Capital Project InVEST Urban Cooling Model to map a heat mitigation index for the study area. The heat mitigation index models the cooling capacity of each pixel and the effect of green spaces on the pixel’s ability to mitigate heat. Results show which communities are most vulnerable to impacts of increased urban heat. The team also assessed alternative tree canopy scenarios with the InVEST model to better understand the effectiveness of potential heat mitigation strategies. Increasing tree canopy cover by 25% in urbanized land cover types was found to reduce mean temperature across the study area by 0.87°F. Combining these results with work from Groundwork’s Climate Safe Neighborhoods program highlights the relationship between extreme heat and historical-race based housing practices in the region. Results provide partners at Groundwork with refined methodologies to support future education and outreach, as well as increase future capacity to NASA Earth observations.

Celeste Gambino↗

NASA NTRSDate not supplied

Time Resolved Visualization of A Liquid Jet in an RDE Using MHz Rate Diesel PLIF

The characterization of the dynamic response of liquid jets to transient detonation wave passage is critical for optimization and modeling of liquid fueled rotating detonation combustors. In this work, a rotating detonation combustor (RDC) is operated on hydrogen and air to sustain stable detonation waves that acts as a detonation driver and interact in a one-way coupled manner with a single liquid fuel jet that propagates into the combustion chamber with cycle periods of ~ 250 μs. Diesel is utilized as a realistic fuel surrogate with higher aromatic compounds to enable fluorescence excitation, using the 355 nm third harmonic output of a burst-mode Nd:YAG laser, imaged at rates up to 1 MHz. By optimizing the technique to accommodate orders of magnitude variations in the fuel density throughout the injection process, the PLIF data enable quantitative measurements including the refill time, the relative recovery between liquid and gaseous jets and jet trajectory across various momentum flux ratio. As the passage of the detonation wave imparts significant changes in the momentum flux ratio, the qualitative liquid break-up process and spatial distribution varies significantly in time. As the injection system recovers ~ 70% of the cycle period and return to a quasi-steady position and allow comparisons with theoretical jet trajectories. These data, enabled by ultra-high-speed PLIF imaging, represent some of the first detailed measurements for quantifying the dynamic response and recovery of liquid jets exposed to periodic detonations in an operating RDC.

Propulsion↗

NASA NTRSDate not supplied

Time Resolved Visualization of Liquid Jet Interaction With H2-Air Rotating Detonations Using MHz Rate Diesel PLIF

The characterization of the dynamic response of liquid jets to transient detonation wave passage is critical for optimization and modeling of liquid fueled rotating detonation combustors. In this work, a rotating detonation combustor (RDC) is operated on hydrogen and air to sustain stable detonation waves that acts as a detonation driver and interact in a one-way coupled manner with a single liquid fuel jet that propagates into the combustion chamber with cycle periods of ~ 250 μs. Diesel is utilized as a realistic fuel surrogate with higher aromatic compounds to enable fluorescence excitation, using the 355 nm thirdharmonic output of a burst-mode Nd:YAG laser, imaged at rates up to 1 MHz. By optimizing the technique to accommodate orders of magnitude variations in the fuel density throughout the injection process, the PLIF data enable quantitative measurements including the refill time, the relative recovery between liquid and gaseous jets and jet trajectory across various momentum flux ratio. As the passage of the detonation wave imparts significant changes in the momentum flux ratio, the qualitative liquid break-up process and spatial distribution varies significantly in time. As the injection system recovers ~ 70% of the cycle period and return to a quasi-steady position and allow comparisons with theoretical jet trajectories. These data, enabled by ultra-high-speed PLIF imaging, represent some of the first detailed measurements for quantifying the dynamic response and recovery of liquid jets exposed to periodic detonations in an operating RDC.

Propulsion↗

NASA NTRSDate not supplied

Propulsive trajectory optimization to minimize surface contamination

MOTIVATION: We present an optimization technique for propulsive vehicles that autonomously minimizes contamination during surface approach and landing. In addition to short-range hoppers, the optimization technique is also fully applicable to traditional orbit-to-surface landers. This study addresses scenarios where surface alterations from propulsion events are counterproductive or hazardous to the mission objectives. This is of immediate interest for landers (whether human or robotic), that may rely on pristine soils collected in the immediate vicinity of landing sites to accomplish science investigations, mining, or ISRU surface operations. Such missions are averse to various surface-plume interactions such as thermal scoring, physical agitation, and contamination. The capability can be applied with minimal impact to the baseline mission concept. METHODS: Optimization algorithms have been developed to calculate descent trajectories and maneuvers, thrust magnitude, and attitude for various mission cases. These parameters are determined as an optimal solution when minimizing either fuel consumption, contamination deposited at the landing site, or some weighted combination of both. Among constraints imposed on the solution, we examined pitch rate, vertical takeoff and vertical landing (VTVL) requirements, size of the contamination zone, and minimum ground clearance during flight. This tool provides unique, non-intuitive solutions and can be a valuable resource for mission planners. RESULTS: A variety of agile trajectory solutions were obtained, each yielding different reductions in landing site contamination and corresponding to only modest increases in fuel consumption. Several optimal trajectories were obtained by varying the contamination weight in the fitness function. As expected, when the contamination weight is zero, the trajectory appears close to parabolic since the optimization scheme only attempts to minimize for fuel utilization, yielding essentially, the expected ballistic trajectory. Notably for contamination weights greater than zero, trajectory inflections are observed in the descent phase, which manifests as hovering or additional, mini “pseudo hops” before the final touchdown. A trajectory inflection is characterized by arresting the majority of the spacecraft vertical velocity component at a coordinate outside of the landing target, and without violating ground clearance constraints. FUTURE WORK: Our optimization technique is ready for laboratory or field demonstrations to validate the sophisticated maneuvering solutions obtained for fuel optimization and surface preservation. An appropriate testbed would validate the optimal guidance algorithms, the navigation system, and sensor suite by emulating vehicle flight in closed loop robotic tests. Critically, these algorithms could then be ported to flight software for implementation.

surface contamination↗

NASA NTRS1984

Recovery of burner acoustic source structure from far-field sound spectra

A method is presented that permits the thermal-acoustic efficiency spectrum in a long turbulent burner to be recovered from the corresponding far-field sound spectrum. An acoustic source/propagation model is used based on the perturbation solution of the equations describing the unsteady one-dimensional flow of an inviscid ideal gas with a distributed heat source. The technique is applied to a long cylindrical hydrogen-flame burner operating over power levels of 4.5-22.3 kW. The results show that the thermal-acoustic efficiency at a given frequency, defined as the fraction of the total burner power converted to acoustic energy at that frequency, is rather insensitive to burner power, having a maximum value on the order of 10 to the -4th at 150 Hz and rolling off steeply with increasing frequency. Evidence is presented that acoustic agitation of the flame at low frequencies enhances the mixing of the unburned fuel and air with the hot products of combustion. The paper establishes the potential of the technique as a useful tool for characterizing the acoustic source structure in any burner, such as a gas turbine combustor, for which a reasonable acoustic propagation model can be postulated.

Mahan, J. R.↗

NASA NTRS1993

Droplet-turbulence interactions in sprays exposed to supercritical environmental conditions

The goal of this research was to experimentally characterize the behavior of droplets in vaporizing sprays under conditions typical of those encountered in high pressure combustion systems such as liquid fueled rocket engines. Of particular interest are measurements of droplet drag and lift, droplet dispersion, droplet heating, and droplet vaporization under both subcritical and supercritical conditions. A summary of the major accomplishments achieved during the period from June 1990 through June 1993, a brief description and status report on five research areas, which were directly or indirectly supported by this grant, and a list of publications and personnel associated with this research is included.

Santavicca, Domenic A.↗

NASA NTRS2003

Microfabricated Liquid Rocket Motors

Under NASA Glenn Research Center sponsorship, MIT has developed the concept of micromachined, bipropellant, liquid rocket engines. This is potentially a breakthrough technology changing the cost-performance tradeoffs for small propulsion systems, enabling new applications, and redefining the meaning of the term low-cost-access-to-space. With this NASA support, a liquid-cooled, gaseous propellant version of the thrust chamber and nozzle was designed, built, and tested as a first step. DARPA is currently funding MIT to demonstrate turbopumps and controls. The work performed herein was the second year of a proposed three-year effort to develop the technology and demonstrate very high power density, regeneratively cooled, liquid bipropellant rocket engine thrust chamber and nozzles. When combined with the DARPA turbopumps and controls, this work would enable the design and demonstration of a complete rocket propulsion system. The original MIT-NASA concept used liquid oxygen-ethanol propellants. The military applications important to DARPA imply that storable liquid propellants are needed. Thus, MIT examined various storable propellant combinations including N2O4 and hydrazine, and H2O2 and various hydrocarbons. The latter are preferred since they do not have the toxicity of N2O4 and hydrazine. In reflection of the newfound interest in H2O2, it is once again in production and available commercially. A critical issue for the microrocket engine concept is cooling of the walls in a regenerative design. This is even more important at microscale than for large engines due to cube-square scaling considerations. Furthermore, the coolant behavior of rocket propellants has not been characterized at microscale. Therefore, MIT designed and constructed an apparatus expressly for this purpose. The report details measurements of two candidate microrocket fuels, JP-7 and JP-10.

Epstein, Alan H.↗

NASA NTRS1954

Experimental Determination of Linear Dynamics of Two-Spool Turbojet Engines

Transfer functions descriptive of the response of most engine variables were determined from transient data that were obtained from approximate step inputs in fuel flow and in exhaust-nozzle area. The speed responses of both spools to fuel flow and to turbine-inlet temperature appeared as identical first-order lags. Response to exhaust-nozzle area was characterized by a first-order lag response of the outer-spool speed, accompanied by virtually no change in inner-spool speed.

Novik, David↗