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At least 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Recommended Crew Systems Capabilities for a 30 to 90 Day Four-Crew Lunar Surface Mission

For any given surface mission duration, there exists a set of surface capabilities necessary to sustain a four-person crew for the length of time the crew is on the surface. Identifying these capabilities is important to mission planners and spacecraft design engineers responsible for lunar surface habitation elements, who need to know which systems and capabilities must be prioritized for the vehicles they are designing. However, simply identifying the surface capabilities does not necessarily equal a viable surface configuration. These capabilities may or may not satisfy stakeholder expectations or even permit completion of mission objectives. Lunar lander cargo mass constraints are a powerful forcing function that requires significant efforts to reduce the mass of all surface elements, including those supporting crew habitation. The required capabilities do increase as a function of mission duration, but it is a nonlinear growth. There are key points in mission duration, beyond which certain additional capabilities are needed. This forms bands of capability between each point. Within a given band, the same functional capabilities are needed, and the only scaling is that associated with daily consumables. The next higher band requires additional capabilities. This paper will focus on those capabilities needed for missions of 30-90 days. It should be noted that there are some uncertainties interpreting NASA requirements that require capabilities after a certain number of days, as typically this refers to number of days in space, not number of days on the lunar surface. Depending on transportation architecture, the number of days between Earth launch and lunar surface landing may vary, therefore impacting when certain required capabilities are implemented. However, activities to prepare humans for missions to Mars and commercial interest in lunar development may drive progressive increases in surface duration. Missions up to 30 days are expected to follow the initial 6.5-day surface missions and there may be commercial interest in extending to even greater durations. This paper will examine the potential case of a 30-90-day surface mission duration with respect to the systems necessary to support crew living and working on the lunar surface.

Habitat↗
NASA NTRSDate not supplied

Liftoff Environments

A Launch Environment, or Liftoff Environment, fundamentally represents a source of loading during the hold-down or liftoff phase of a rocket launch. A launch vehicle is subjected to transient and oscillatory pressure loading on its exterior surfaces induced by engines and boosters during vehicle hold-down and liftoff. These pressure loads are a principal source of structural vibration which may result in the malfunction of vehicle components or the fatigue of exterior skin panels or component brackets. Additionally, the pressure loading gives rise to sound pressure levels inside the crew cabin which may impact the crew's health, safety, or ability to communicate. Liftoff Environments can have a significant impact on the launch vehicle or surrounding structures. For SLS, there are eight primary environments supported by the Liftoff Environments team. These are Liftoff Acoustics (LOA), Hold-down Acoustics (HDA), RS-25 Nozzle Flow Transient (NFT) Acoustics, Booster Ignition Overpressure (IOP) & Duct Overpressure (DOP), Igniter Shock, Core Stage Engine (CSE) Overpressure, Excess Hydrogen (H2) Pop, and Infrasonic Acoustics. These can be split generally into acoustics, which are sustained stationary events, and transients, which are short duration nonstationary events.

Matthew J Casiano↗
NASA NTRSDate not supplied

Study of the Thermal Physical Properties of Insulating Materials in A Titan Environment

Dragonfly is a planned rotorcraft mission that NASA/APL will be sending to Saturn’s moon Titan to study its chemistry. The Dragonfly Mass Spectrometer (DraMS) is a mass spectrometer that is being developed to identify different kinds of organic material that comprises Titan’s surface. The Cryogenic engineering team on DraMS is developing and interface between the room temperature rotorcraft body and the near-cryogenic temperatures of an onboard sample chamber, while minimizing the thermal leak to the Titan environment. As part of design process, the Cryogenic team has built a conductivity test rig for testing the thermophysical properties of candidate insulators, which includes Rohacell 31HF and hollowed 3D printed PEEK structures, at simulated Titan environmental conditions.

Thermal Conductivity↗
NASA NTRS2014

Possible Mechanisms for Turbofan Engine Ice Crystal Icing at High Altitude

A thermodynamic model is presented to describe possible mechanisms of ice formation on unheated surfaces inside a turbofan engine compression system from fully glaciated ice crystal clouds often formed at high altitude near deep convective weather systems. It is shown from the analysis that generally there could be two distinct types of ice formation: (1) when the "surface freezing fraction" is in the range of 0 to 1, dominated by the freezing of water melt from fully or partially melted ice crystals, the ice structure is formed from accretion with strong adhesion to the surface, and (2) when the "surface melting fraction" is the range of 0 to 1, dominated by the further melting of ice crystals, the ice structure is formed from accumulation of un-melted ice crystals with relatively weak bonding to the surface. The model captures important qualitative trends of the fundamental ice-crystal icing phenomenon reported earlier1,2 from the research collaboration work by NASA and the National Research Council (NRC) of Canada. Further, preliminary analysis of test data from the 2013 full scale turbofan engine ice crystal icing test3 conducted in the NASA Glenn Propulsion Systems Laboratory (PSL) has also suggested that (1) both types of ice formation occurred during the test, and (2) the model has captured some important qualitative trend of turning on (or off) the ice crystal ice formation process in the tested engine low pressure compressor (LPC) targeted area under different icing conditions that ultimately would lead to (or suppress) an engine core roll back (RB) event.

Crystal Icing↗
NASA NTRS2016

Possible Mechanisms for Turbofan Engine Ice Crystal Icing at High Altitude

A thermodynamic model is presented to describe possible mechanisms of ice formation on unheated surfaces inside a turbofan engine compression system from fully glaciated ice crystal clouds often formed at high altitude near deep convective weather systems. It is shown from the analysis that generally there could be two distinct types of ice formation: (1) when the "surface freezing fraction" is in the range of 0 to 1, dominated by the freezing of water melt from fully or partially melted ice crystals, the ice structure is formed from accretion with strong adhesion to the surface, and (2) when the "surface melting fraction" is the range of 0 to 1, dominated by the further melting of ice crystals, the ice structure is formed from accumulation of un-melted ice crystals with relatively weak bonding to the surface. The model captures important qualitative trends of the fundamental ice-crystal icing phenomenon reported earlier (Refs. 1 and 2) from the research collaboration work by NASA and the National Research Council (NRC) of Canada. Further, preliminary analysis of test data from the 2013 full scale turbofan engine ice crystal icing test (Ref. 3) conducted in the NASA Glenn Propulsion Systems Laboratory (PSL) has also suggested that (1) both types of ice formation occurred during the test, and (2) the model has captured some important qualitative trend of turning on (or off) the ice crystal ice formation process in the tested engine low pressure compressor (LPC) targeted area under different icing conditions that ultimately would lead to (or suppress) an engine core roll back (RB) event.

Crystal Icing↗
NASA NTRS1989

Low-speed, high-lift aerodynamic characteristics of slender, hypersonic accelerator-type configurations

Two investigations were conducted in the Langley 14 by 22 Foot Subsonic Tunnel to determine the low-speed aerodynamic characteristics of a generic hypersonic accelerator-type configuration. The model was a delta wing configuration incorporating a conical forebody, a simulated wrap-around engine package, and a truncated conical aftbody. Six-component force and moment data were obtained over a range of attack from -4 to 30 degrees and for a sideslip range of + or - 20 degrees. In addition to tests of the basic configuration, component build-up tests were conducted; and the effects of power, forebody nose geometry, canard surfaces, fuselage strakes, and engines on the lower surface alone were also determined. Control power available from deflections of wing flaps and aftbody flaps was also investigated and found to be significantly increased during power-on conditions. Large yawing moments resulted from asymmetric flow fields exhibited by the forebody as revealed by both surface pressure data and flow visualization. Increasing nose bluntness reduced the yawing-moment asymmetry, and the addition of a canard eliminated the yawing-moment asymmetry.

Gatlin, Gregory M.↗
NASA NTRS2015

Data-Driven Surface Traversability Analysis for Mars 2020 Landing Site Selection

The objective of this paper is three-fold: 1) to describe the engineering challenges in the surface mobility of the Mars 2020 Rover mission that are considered in the landing site selection processs, 2) to introduce new automated traversability analysis capabilities, and 3) to present the preliminary analysis results for top candidate landing sites. The analysis capabilities presented in this paper include automated terrain classification, automated rock detection, digital elevation model (DEM) generation, and multi-ROI (region of interest) route planning. These analysis capabilities enable to fully utilize the vast volume of high-resolution orbiter imagery, quantitatively evaluate surface mobility requirements for each candidate site, and reject subjectivity in the comparison between sites in terms of engineering considerations. The analysis results supported the discussion in the Second Landing Site Workshop held in August 2015, which resulted in selecting eight candidate sites that will be considered in the third workshop.

Ono, Masahiro↗
NASA NTRS1980

Space shuttle status

Technology problems relating to the shuttle thermal protection system and the auxiliary power units for flight control of engine gimbaling and aero surfaces are examined as well as yet unresolved difficulties associated with the main engine, external tank doors, and the payload bay doors. Concepts for launch crew escape and for flight inspection are discussed, and the orbital flight test drivers are identified. The majority of OFT objectives can be accomplished in a success oriented four flight program. Carry over will be accommodated in the operational phase with a mini-DFI package. Primary shuttle systems requiring additional testings are (1) aerodynamics (forward CG); (2) thermal control system (high beta); (3) payload deployment and retrieval system (retrieval, a maximum weight, and proximity operations); (4) SSME throttling to 109%; and (5) SSME nozzle entry aerodynamic heating (up elevon/body flap position at hypersonic velocities).

Slayton, D. K.↗
NASA NTRS1990

A note on an acoustic response during an engine nacelle flight experiment

During a flight test study of the noise effects on laminar flow on the outside surface of a simulated engine nacelle, an intense acoustic response was observed. The aircraft speed at which this signal occurred and the frequency content of the signal fell within the test conditions of the experiment and had to be eliminated prior to continuing. The signal was identified as an aerodynamic excitation of an acoustic mode in the simulated by-pass duct of the nacelle. By modifying the trailing edges of the support struts of the nacelle, the aerodynamic excitation was changed enough to eliminate the resonant response of the offending duct modes, eliminating the unwanted acoustic problem.

Schoenster, James A.↗
NASA NTRS1977

Integrated gas turbine engine-nacelle

A nacelle for use with a gas turbine engine is presented. An integral webbed structure resembling a spoked wheel for rigidly interconnecting the nacelle and engine, provides lightweight support. The inner surface of the nacelle defines the outer limits of the engine motive fluid flow annulus while the outer surface of the nacelle defines a streamlined envelope for the engine.

Adamson, A. P.↗
NASA NTRS2011

Further Analysis on the Mystery of the Surveyor III Dust Deposits

The Apollo 12 lunar module (LM) landing near the Surveyor 1lI spacecraft at the end of 1969 has remained the primary experimental verification of the predicted physics of plume ejecta effects from a rocket engine interacting with the surface of the moon. This was made possible by the return of the Surveyor 1lI camera housing by the Apollo 12 astronauts, allowing detailed analysis of the composition of dust deposited by the Apollo 12 LM plume. It was soon realized after the initial analysis of the camera housing that the LM plume tended to remove more dust than it had deposited. In the present study, coupons from the camera housing were reexamined by a KSC research team using SEM/EDS and XPS analysis. In addition, plume effects recorded in landing videos from each Apollo mission have been studied for possible clues. Several likely scenarios are proposed to explain the Surveyor III dust observations. These include electrostatic attraction of the dust to the surface of the Surveyor as a result of electrostatic charging of the jet gas exiting the engine nozzle during descent; dust blown by the Apollo 12 LM fly-by while on its descent trajectory; dust ejected from the lunar surface due to gas forced into the soil by the Surveyor 1lI rocket nozzle, based on Darcy's law; and mechanical movement of dust during the Surveyor landing. Even though an absolute answer is not possible based on available data and theory, various computational models are employed to estimate the feasibility of each of these proposed mechanisms. Scenarios are then discussed which combine multiple mechanisms to produce results consistent with observations.

Metzger, Philip↗
NASA NTRS1975

Effect of empennage interference on single-engine afterbody/nozzle drag

An investigation has been conducted in the Langley 16-foot transonic tunnel to determine the effects of empennage surfaces on single-engine afterbody and nozzle drag at Mach numbers up to 1.20. Empennage interference drag was obtained by using experimental values of afterbody/nozzle drag and theoretical values of empennage drag. Three methods for minimizing adverse empennage interference effects were investigated. Results show that empennage interference effects were always detrimental but could be reduced by proper location of tail surfaces. Afterbody contouring and addition of 'contour' bumps were generally ineffective in reducing aft-end drag for Mach numbers less than 0.95.

Berrier, B. L.↗
NASA NTRS2013

Background Oriented Schlieren Implementation in a Jet-Surface Interaction Test

Many current and future aircraft designs rely on the wing or other aircraft surfaces to shield the engine noise from observers on the ground. However the available data regarding how a planar surface interacts with a jet to shield and/or enhance the jet noise are currently limited. Therefore, the Jet-Surface Interaction Tests supported by NASA's Fundamental Aeronautics Program's Fixed Wing Project were undertaken to supply experimental data covering a wide range of surface geometries and positions interacting with high-speed jet flows in order to support the development of noise prediction methods. Phase 1 of the Test was conducted in the Aero-Acoustic Propulsion Laboratory at NASA Glenn Research Center and consisted of validating noise prediction schemes for a round nozzle interacting with a planar surface. Phased array data and far-field acoustic data were collected for both the shielded and reflected sides of the surface. Phase 1 results showed that the broadband shock noise was greatly reduced by the surface when the jet was operated at the over-expanded condition, however, it was unclear whether this reduction was due a change in the shock cell structure by the surface. In the present study, Background Oriented Schlieren is implemented in Phase 2 of the Jet-Surface Interaction Tests to investigate whether the planar surface interacts with the high-speed jet ow to change the shock cell structure. Background Oriented Schlieren data are acquired for under-expanded, ideally-expanded, and over-expanded ow regimes for multiple axial and radial positions of the surface at three different plate lengths. These data are analyzed with far-field noise measurements to relate the shock cell structure to the broadband shock noise produced by a jet near a surface.

Clem, Michelle M.↗
NASA NTRSDate not supplied

Design Study of Surface to Surface Laser Power Beaming on the Moon

Engineering design of a near-term laser surface-to-surface power beaming station to transmit power to users in shadowed regions or during the night. Input system requirements were to be able to provide 300 W of continuous usable power to users including landers and rovers, at distances up to 10 km, with a total system landed mass under 625 kg. Due to surface irregularities and the close horizon of the moon, to achieve 10-km transmission the laser must be elevated above the surface. The Vertical Solar Array Technology (VSAT) program is developing a solar array mounted on a 10-m tall mast, intended to fit on a Commercial Lunar Payload Services (CLPS) lander, with target readiness date of 2028. At an optimum location near the south pole, the elevated array produces power for a majority of the lunar day.

Vertical Solar Array Technologoy (VSAT)↗
NASA NTRS1981

Modification of spacecraft potentials by plasma emission

The ion engine operations on Applied Technology Satellite 6 (ATS-6) altered the charge state of the spacecraft, changing the spacecraft surface potentials with respect to the distant plasmas. Plasma emission in quiet environments (plasma temperatures below 1 keV) caused the spacecraft potential to shift from a few volts positive to a few volts negative. A net ion current is emitted in such cases. The emission of a plasma or beam in energetic environments (plasma temperatures in the 5-10-keV range) in sunlight caused larger changes. Typical equilibrium potentials for ATS-6 in such environments were on the order of a hundred volts negative, with variations in potential across the spacecraft surface of comparable magnitude. Engine operations under such conditions raised the mainframe potential to near zero volts, and discharged the differential potentials on the dielectric surfaces. Plasma emission by the plasma bridge was an effective method of discharging kilovolt potentials in eclipse.

Olsen, R. C.↗
NASA NTRS1974

Theoretical surface velocity distributions on acoustic splitter geometries for an engine inlet

The potential-flow velocity distributions on several splitter geometries in an engine inlet and their variation with different splitter leading-edge shapes and distances from the inlet highlight were analyzed. The velocity distributions on the inner and outer surfaces of the splitters are presented for low-speed and cruise conditions. At zero incidence angle, the splitter with the 4-to-1 elliptical leading edge had lower peak velocities and velocity gradients than the splitter with the 2-to-1 elliptical leading edge. The velocity gradients decreased as the distance from the inlet highlight to the splitter leading edge was increased. For a given distance, the peak velocity on the splitter inner surface increased with increasing inlet incidence angle. At an incidence angle of 50 deg, the velocity level and gradients on the inner surface of the splitter in the forward position were sufficiently severe to suggest local separation.

Albers, J. A.↗
NASA NTRS1998

Experimental Techniques Verified for Determining Yield and Flow Surfaces

Structural components in aircraft engines are subjected to multiaxial loads when in service. For such components, life prediction methodologies are dependent on the accuracy of the constitutive models that determine the elastic and inelastic portions of a loading cycle. A threshold surface (such as a yield surface) is customarily used to differentiate between reversible and irreversible flow. For elastoplastic materials, a yield surface can be used to delimit the elastic region in a given stress space. The concept of a yield surface is central to the mathematical formulation of a classical plasticity theory, but at elevated temperatures, material response can be highly time dependent. Thus, viscoplastic theories have been developed to account for this time dependency. Since the key to many of these theories is experimental validation, the objective of this work (refs. 1 and 2) at the NASA Lewis Research Center was to verify that current laboratory techniques and equipment are sufficient to determine flow surfaces at elevated temperatures. By probing many times in the axial-torsional stress space, we could define the yield and flow surfaces. A small offset definition of yield (10 me) was used to delineate the boundary between reversible and irreversible behavior so that the material state remained essentially unchanged and multiple probes could be done on the same specimen. The strain was measured with an off-the-shelf multiaxial extensometer that could measure the axial and torsional strains over a wide range of temperatures. The accuracy and resolution of this extensometer was verified by comparing its data with strain gauge data at room temperature. The extensometer was found to have sufficient resolution for these experiments. In addition, the amount of crosstalk (i.e., the accumulation of apparent strain in one direction when strain in the other direction is applied) was found to be negligible. Tubular specimens were induction heated to determine the flow surfaces at elevated temperatures. The heating system induced a large amount of noise in the data. By reducing thermal fluctuations and using appropriate data averaging schemes, we could render the noise inconsequential. Thus, accurate and reproducible flow surfaces (see the figure) could be obtained.

Lerch, Brad A.↗