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NASA NTRSDate not supplied

NASA’s Space Launch System: New Launch Capability for Artemis Lunar and Deep Space Science Missions

With stacking and integration of the initial Block 1 Space Launch System (SLS) expected to begin in 2020, NASA’s powerful new launch vehicle is ready to take center stage in the agency’s Artemis program to return astronauts to the Moon. Combining the highest launch thrust and largest payload capacity ever developed, SLS also enables a new generation of high-C3 science missions to destinations such as the gas and ice giants, the Kuiper Belt, and even beyond the solar system. Block 1 is only the beginning, as the vehicle has a planned evolution path to progressively more powerful variants. In addition to these block upgrades providing increased lift capability, the vehicle can be configured to fly in crew configuration with the Orion spacecraft or in cargo configuration with payload fairings for launching science mission or large infrastructure, providing a flexible launch option. For Artemis I, the first SLS flight, the Block 1 vehicle in the crew configuration will send an uncrewed Orion spacecraft to lunar orbit for a thorough systems checkout before the crewed Artemis II flight. The Block 1 vehicle uses a proven propulsion system consisting of solid rocket boosters and RS-25 engines to lift more than 27 metric tons [t] to trans-lunar injection (TLI). In its cargo configuration, Block 1 can be fitted with a 5 m payload fairing. The second variant, Block 1B, uses a more powerful upper stage to increase payload mass to TLI to 38-42 t, depending on crew or cargo configuration. In the crew configuration, a co-manifested payload of up to 10 t can ride along in the Universal Stage Adapter (USA), which has as much volume for payloads as a 5 m-class payload fairing. The Block 2 evolved variant will lift 43-46 t to TLI, depending on crew or cargo configuration. The Block 1B and Block 2 vehicles can be outfitted with an 8.4 m-diameter payload fairing, available in 19.1 m and 27.4 m lengths, providing unprecedented volume for payloads. Larger-diameter 10 m fairings may also be an option in the future on the Block 2 vehicle. The unrivalled mass, volume and high-energy launches of SLS can provide significant mission flexibility for payloads and/or additional upper stages to open trade space for a new generation of exploration missions. SLS was designed to meet requirements for launching large-volume infrastructure as outlined in numerous studies of missions to cislunar space or Mars. Mission concept studies from the science community also point toward new possibilities enabled by SLS. Probes with more robust science packages can be sent to the gas giants. Dual spacecraft can be manifested for missions to Uranus and Neptune. Additional third or fourth payload stages can be encapsulated in the payload fairings to achieve missions to the Kuiper Belt or beyond. In addition, the large volume can be used to design and deploy wide-aperture mirrors on future space telescopes and to enable nuclear-thermal propulsion missions. At AIAA Ascend, the SLS Program will provide technical information on vehicle capabilities as well as descriptions of ongoing discussions with mission planners for utilizing the vehicle for an array of deep space missions.

Stephen Creech↗

NASA NTRS2017

The Effectiveness of Shrouding on Reducing Meshed Spur Gear Power Loss - Test Results

Gearbox efficiency is reduced at high rotational speeds due to windage drag and viscous effects on rotating, meshed gear components. A goal of NASA aeronautics rotorcraft research is aimed at propulsion technologies that improve efficiency while minimizing vehicle weight. Specifically, reducing power losses to rotorcraft gearboxes would allow gains in areas such as vehicle payload, range, mission type, and fuel consumption. To that end, a gear windage rig has been commissioned at NASA Glenn Research Center to measure windage drag on gears and to test methodologies to mitigate windage power losses. One method used in rotorcraft gearbox design attempts to reduce gear windage power loss by utilizing close clearance walls to enclose the gears in both the axial and radial directions. The close clearance shrouds result in reduced drag on the gear teeth, and reduced power loss. For meshed spur gears, the shrouding takes the form of metal side plates and circumferential metal sectors. Variably positioned axial and radial shrouds are incorporated in the NASA rig to study the effect of shroud clearance on gearbox power loss. A number of researchers have given experimental and analytical results for single spur gears, with and without shrouding. Shrouded meshed spur gear test results are sparse in the literature. Windage tests were run at NASA Glenn using meshed spur gears at four shroud configurations: unshrouded, shrouded (max. axial, max radial), and two intermediate shrouding conditions. Results are compared to available meshed spur gear power loss data analyses as well as single spur gear data/analyses. Recommendations are made for future work.

power loss↗

NASA NTRS2017

The Effectiveness of Shrouding on Reducing Meshed Spur Gear Power Loss Test Results

Gearbox efficiency is reduced at high rotational speeds due to windage drag and viscous effects on rotating, meshed gear components. A goal of NASA aeronautics rotorcraft research is aimed at propulsion technologies that improve efficiency while minimizing vehicle weight. Specifically, reducing power losses to rotorcraft gearboxes would allow gains in areas such as vehicle payload, range, mission type, and fuel consumption. To that end, a gear windage rig has been commissioned at NASA Glenn Research Center to measure windage drag on gears and to test methodologies to mitigate windage power losses. One method used in rotorcraft gearbox design attempts to reduce gear windage power loss by utilizing close clearance walls to enclose the gears in both the axial and radial directions. The close clearance shrouds result in reduced drag on the gear teeth and reduced power loss. For meshed spur gears, the shrouding takes the form of metal side plates and circumferential metal sectors. Variably positioned axial and radial shrouds are incorporated in the NASA rig to study the effect of shroud clearance on gearbox power loss. A number of researchers have given experimental and analytical results for single spur gears, with and without shrouding. Shrouded meshed spur gear test results are sparse in the literature. Windage tests were run at NASA Glenn using meshed spur gears at four shroud configurations: unshrouded, shrouded (max. axial, max. radial), and two intermediate shrouding conditions. Results are compared to available meshed spur gear power loss data analyses as well as single spur gear data analyses.

windage↗

NASA NTRS2011

Computational Assessment of the Aerodynamic Performance of a Variable-Speed Power Turbine for Large Civil Tilt-Rotor Application

The main rotors of the NASA Large Civil Tilt-Rotor notional vehicle operate over a wide speed-range, from 100% at take-off to 54% at cruise. The variable-speed power turbine offers one approach by which to effect this speed variation. Key aero-challenges include high work factors at cruise and wide (40 to 60 deg.) incidence variations in blade and vane rows over the speed range. The turbine design approach must optimize cruise efficiency and minimize off-design penalties at take-off. The accuracy of the off-design incidence loss model is therefore critical to the turbine design. In this effort, 3-D computational analyses are used to assess the variation of turbine efficiency with speed change. The conceptual design of a 4-stage variable-speed power turbine for the Large Civil Tilt-Rotor application is first established at the meanline level. The design of 2-D airfoil sections and resulting 3-D blade and vane rows is documented. Three-dimensional Reynolds Averaged Navier-Stokes computations are used to assess the design and off-design performance of an embedded 1.5-stage portion-Rotor 1, Stator 2, and Rotor 2-of the turbine. The 3-D computational results yield the same efficiency versus speed trends predicted by meanline analyses, supporting the design choice to execute the turbine design at the cruise operating speed.

Welch, Gerard E.↗

NASA NTRS1992

Phase change energy storage for solar dynamic power systems

This paper presents the results of a transient computer simulation that was developed to study phase change energy storage techniques for Space Station Freedom (SSF) solar dynamic (SD) power systems. Such SD systems may be used in future growth SSF configurations. Two solar dynamic options are considered in this paper: Brayton and Rankine. Model elements consist of a single node receiver and concentrator, and takes into account overall heat engine efficiency and power distribution characteristics. The simulation not only computes the energy stored in the receiver phase change material (PCM), but also the amount of the PCM required for various combinations of load demands and power system mission constraints. For a solar dynamic power system in low earth orbit, the amount of stored PCM energy is calculated by balancing the solar energy input and the energy consumed by the loads corrected by an overall system efficiency. The model assumes an average 75 kW SD power system load profile which is connected to user loads via dedicated power distribution channels. The model then calculates the stored energy in the receiver and subsequently estimates the quantity of PCM necessary to meet peaking and contingency requirements. The model can also be used to conduct trade studies on the performance of SD power systems using different storage materials.

Chiaramonte, F. P.↗

NASA NTRSDate not supplied

Onboard Autonomous Trajectory Planning for Mars Power Descent

In recent years, there has been an increasing interest in space-qualified processors such as multi-core central processing units and graphics processing units that can withstand the adverse effects of space radiation. These processors can allow parallel programming to perform tasks that typically demand high computational power. One can study guidance schemes that can take advantage of these currently developing processors and provide more robust guidance. Software for Multi-model Autonomous Real-time Trajectories (SMART) guidance can identify robust trajectories by running an onboard Monte Carlo analysis. SMART guidance can take advantage of knowledge updates obtained from the onboard sensors, allowing it to consider the off-nominal cases that it would not typically encounter during the offline trajectory analysis. This work uses the SMART guidance for the powered divert at Mars simulation in Program to Optimize and Simulated Trajectories- II.

Pardha Sai Chadalavada↗

NASA NTRSDate not supplied

Onboard Autonomous Trajectory Planning for Mars Power Descent

Autonomous Planning↗

NASA NTRS2013

Raster-Based Approach to Solar Pressure Modeling

An algorithm has been developed to take advantage of the graphics processing hardware in modern computers to efficiently compute high-fidelity solar pressure forces and torques on spacecraft, taking into account the possibility of self-shading due to the articulation of spacecraft components such as solar arrays. The process is easily extended to compute other results that depend on three-dimensional attitude analysis, such as solar array power generation or free molecular flow drag. The impact of photons upon a spacecraft introduces small forces and moments. The magnitude and direction of the forces depend on the material properties of the spacecraft components being illuminated. The parts of the components being lit depends on the orientation of the craft with respect to the Sun, as well as the gimbal angles for any significant moving external parts (solar arrays, typically). Some components may shield others from the Sun. The purpose of this innovation is to enable high-fidelity computation of solar pressure and power generation effects of illuminated portions of spacecraft, taking self-shading from spacecraft attitude and movable components into account. The key idea in this innovation is to compute results dependent upon complicated geometry by using an image to break the problem into thousands or millions of sub-problems with simple geometry, and then the results from the simpler problems are combined to give high-fidelity results for the full geometry. This process is performed by constructing a 3D model of a spacecraft using an appropriate computer language (OpenGL), and running that model on a modern computer's 3D accelerated video processor. This quickly and accurately generates a view of the model (as shown on a computer screen) that takes rotation and articulation of spacecraft components into account. When this view is interpreted as the spacecraft as seen by the Sun, then only the portions of the craft visible in the view are illuminated. The view as shown on the computer screen is composed of up to millions of pixels. Each of those pixels is associated with a small illuminated area of the spacecraft. For each pixel, it is possible to compute its position, angle (surface normal) from the view direction, and the spacecraft material (and therefore, optical coefficients) associated with that area. With this information, the area associated with each pixel can be modeled as a simple flat plate for calculating solar pressure. The vector sum of these individual flat plate models is a high-fidelity approximation of the solar pressure forces and torques on the whole vehicle. In addition to using optical coefficients associated with each spacecraft material to calculate solar pressure, a power generation coefficient is added for computing solar array power generation from the sum of the illuminated areas. Similarly, other area-based calculations, such as free molecular flow drag, are also enabled. Because the model rendering is separated from other calculations, it is relatively easy to add a new model to explore a new vehicle or mission configuration. Adding a new model is performed by adding OpenGL code, but a future version might read a mesh file exported from a computer-aided design (CAD) system to enable very rapid turnaround for new designs

Wright, Theodore W. II↗

NASA NTRS2013

Li-Ion Battery and Supercapacitor Hybrid Design for Long Extravehicular Activities

With the need for long periods of extravehicular activities (EVAs) on the Moon or Mars or a near-asteroid, the need for long-performance batteries has increased significantly. The energy requirements for the EVA suit, as well as surface systems such as rovers, have increased significantly due to the number of applications they need to power at the same time. However, even with the best state-of-the-art Li-ion batteries, it is not possible to power the suit or the rovers for the extended period of performance. Carrying a charging system along with the batteries makes it cumbersome and requires a self-contained power source for the charging system that is usually not possible. An innovative method to charge and use the Li-ion batteries for long periods seems to be necessary and hence, with the advent of the Li-ion supercapacitors, a method has been developed to extend the performance period of the Li-ion power system for future exploration applications. The Li-ion supercapacitors have a working voltage range of 3.8 to 2.5 V, and are different from a traditional supercapacitor that typically has a working voltage of 1 V. The innovation is to use this Li-ion supercapacitor to charge Liion battery systems on an as-needed basis. The supercapacitors are charged using solar arrays and have battery systems of low capacity in parallel to be able to charge any one battery system while they provide power to the application. Supercapacitors can safely take up fast charge since the electrochemical process involved is still based on charge separation rather than the intercalation process seen in Li-ion batteries, thus preventing lithium metal deposition on the anodes. The lack of intercalation and eliminating wear of the supercapacitors allows for them to be charged and discharged safely for a few tens of thousands of cycles. The Li-ion supercapacitors can be charged from the solar cells during the day during an extended EVA. The Liion battery used can be half the capacity required for a nominal EVA. The small Li-ion battery can be divided into two parallel modules with independent charging ports that would allow the supercapacitors to charge one battery while the other is providing power to the rover or suit.

Jeevarajan, Judith↗

NASA NTRS1980

Design study: A 186 kW lightweight diesel aircraft engine

The design of an aircraft engine capable of developing 186 kW shaft power at a 7620 m altitude is described. The 186 kW design takes into account expected new developments in aircraft designs resulting in a reassessment of the power requirements at the cruise mode operation. Based on the results of this analysis a three phase technology development program is projected resulting in production dates of 1985, 1992, and 2000.

Brouwers, A. P.↗

NASA NTRS1934

The calculated effect of trailing-edge flaps on the take-off of flying boats

The results of take-off calculations are given for an application of simple trailing-edge flaps to two hypothetical flying boats, one having medium wing and power loading and consequently considerable excess of thrust over total resistance during the take-off run, the other having high wing and power loading and a very low excess thrust. For these seaplanes the effect of downward flap settings was: (1) to increase the total resistance below the stalling speed, (2) to decrease the get-away speed, (3) to improve the take-off performance of the seaplane having considerable excess thrust, and (4) to hinder the take-off of the seaplane having low excess thrust. It is indicated that flaps would allow a decrease in the high angles of wing setting necessary with most seaplanes, provided that the excess thrust is not too low.

Parkinson, J E↗

NASA NTRS1978

High temperature thermal energy storage, including a discussion of TES integrated into power plants

Storage temperatures of 260 C and above are considered. Basic considerations concerning energy thermal storage are discussed, taking into account general aspects of thermal energy storage, thermal energy storage integrated into power plants, thermal storage techniques and technical considerations, and economic considerations. A description of system concepts is provided, giving attention to a survey of proposed concepts, storage in unpressurized fluids, water storage in pressurized containers, the use of an underground lined cavern for water storage, a submerged thin insulated steel shell under the ocean containing pressurized water, gas passage through solid blocks, a rock bed with liquid heat transport fluid, hollow steel ingots, heat storage in concrete or sand, sand in a fluidized bed, sand poured over pipes, a thermal energy storage heat exchanger, pipes or spheres filled with phase change materials (PCM), macroencapsulated PCM with heat pipe concept for transport fluid, solid PCM removed from heat transfer pipes by moving scrapers, and the direct contact between PCM and transport fluid.

Turner, R. H.↗

NASA NTRS2015

The Role of Margin in Link Design and Optimization

Link analysis is a system engineering process in the design, development, and operation of communication systems and networks. Link models that are mathematical abstractions representing the useful signal power and the undesirable noise and attenuation effects (including weather effects if the signal path transverses through the atmosphere) that are integrated into the link budget calculation that provides the estimates of signal power and noise power at the receiver. Then the link margin is applied which attempts to counteract the fluctuations of the signal and noise power to ensure reliable data delivery from transmitter to receiver. (Link margin is dictated by the link margin policy or requirements.) A simple link budgeting approach assumes link parameters to be deterministic values typically adopted a rule-of-thumb policy of 3 dB link margin. This policy works for most S- and X-band links due to their insensitivity to weather effects. But for higher frequency links like Ka-band, Ku-band, and optical communication links, it is unclear if a 3 dB link margin would guarantee link closure. Statistical link analysis that adopted the 2-sigma or 3-sigma link margin incorporates link uncertainties in the sigma calculation. (The Deep Space Network (DSN) link margin policies are 2-sigma for downlink and 3-sigma for uplink.) The link reliability can therefore be quantified statistically even for higher frequency links. However in the current statistical link analysis approach, link reliability is only expressed as the likelihood of exceeding the signal-to-noise ratio (SNR) threshold that corresponds to a given bit-error-rate (BER) or frame-error-rate (FER) requirement. The method does not provide the true BER or FER estimate of the link with margin, or the required signalto-noise ratio (SNR) that would meet the BER or FER requirement in the statistical sense. In this paper, we perform in-depth analysis on the relationship between BER/FER requirement, operating SNR, and coding performance curve, in the case when the channel coherence time of link fluctuation is comparable or larger than the time duration of a codeword. We compute the "true" SNR design point that would meet the BER/FER requirement by taking into account the fluctuation of signal power and noise power at the receiver, and the shape of the coding performance curve. This analysis yields a number of valuable insights on the design choices of coding scheme and link margin for the reliable data delivery of a communication system - space and ground. We illustrate the aforementioned analysis using a number of standard NASA error-correcting codes.

Cheung, K.↗

NASA NTRS2024

Revisiting Optimal Guidance Solution for Vertical Landing and Take-Off

The vertical landing and take-off are crucial phases of a rocket-powered spacecraft's powered descent and ascent trajectories, during which its motion is confined to a single (vertical) axis. It is well-known that the minimum-propellant guidance solutions for vertical landing and take-off in a vacuum, consist of only one coast and one full-thrust arc and no singular arc exists in these problems. However, an exact solution for the vertical descent problem is not found in the literature. In the present work, exact guidance solutions for the optimal vertical landing and take-off are presented that are valid for arbitrary boundary conditions. The time-based and switching function-based guidance law implementations are discussed and their performance is compared using lunar descent and ascent simulations.

powered descent↗

NASA NTRS1973

Research on future short-haul aircraft at the NASA Langley Research Center.

Some of the current research for improving our short-haul air-transportation system is reviewed. Promising aircraft range from helicopters, turbofan-powered STOL, through mechanical flap reduced take-off and landing (RTOL) concepts. Advanced rotorcraft technology can provide improved passenger, community, operational, and economic acceptability of civil transport helicopters. From wind-tunnel and design studies, techniques are available for achieving low-noise fixed-wing STOL and RTOL through proper engine and airframe design. Agreeable ride qualities, crosswind landing capability, and all-weather terminal operation are also goals of present effort.

Riebe, J. M.↗

NASA NTRS1975

Preliminary measurements of plasma fluctuations in an 8-cm mercury ion thruster

The investigation considered included an observation of the effects on the plasma fluctuations of a variation of several neutralizer parameters. The apparatus used in the investigation and the experimental techniques are discussed, taking into account the thruster, the thruster power supply, questions of thruster operation, transducers, and an analysis of time-varying data. The results presented are related to rms magnitudes, spectral data, and fluctuation correlations.

Serafini, J. S.↗

NASA NTRS1981

Comparison of predicted engine core noise with current and proposed aircraft noise certification requirements

Predicted engine core noise levels for subsonic CTOL aircraft engines are compared with measured total aircraft noise levels and to current and proposed federal noise certification requirements. Comparisons are made at FAR-36 measuring stations and take into consideration both full and cutback power operations at takeoff. The spectral shape used for the prediction of core noise is identified as the spectral envelope, with a peak at 400 Hz which is assumed to be shifted in flight by a Doppler shift in frequency. Preceived noise levels are computed for appropriate engine power settings at desired flight conditions, and reductions in sideline noise levels are made to account for jet and airframe shielding effects. Results indicate that core noise can provide a barrier to the proposed EPA stage 4 and 5 federal noise rules for wide-body aircraft, with the most severe core noise problem occurring at takeoff and sideline measuring stations.

Von Glahn, U. H.↗

NASA NTRS1985

Photographic Screening for Eye Defects

Images of retinas examined for characteristic patterns. Color photographs of retinas taken. Proper alinement of eye obtained by asking subject to gaze at light-emitting diode. "Red-eye" patterns in resulting color photographs examined by trained observers for signs of ocular defects. System used to check power of contact lenses and eyeglasses by taking photographs with these items in place.

Richardson, J.↗