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At least 721 records · Page 40

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NASA NTRS2018

An Innovative Control Allocation Scheme to Address Reaction Thruster Interactions on a 3U CubeSat

A wealth of literature exists on control allocation algorithms for over-actuated air vehicles, launch vehicles, and spacecraft's. Most of these algorithms focus primarily on minimizing some objective function such as command tracking error and/or control effector usage. Linear allocators (pseudo inverses) are usually the conventional choice due to their simplicity and the ability to achieve a significant portion of the theoretical moment/impulse space. Generally, it is assumed that there exists minimal interaction effects between control effectors. In fact, very few studies address the problem of control effector interactions in the context of control allocation, especially for small spacecraft's with a reaction control system (RCS). This paper presents a CubeSat RCS design with a four thruster tetrahedral layout such that when two or more thrusters re, the resultant impulse differs noticeably compared to the sum of the contributions from individual thruster rings. This undesirable effect is caused by the design of the propellant tank and regulator. To mitigate this issue, an innovative modified pseudo inverse (MPI) control allocation algorithm was developed that adjusts the pseudo inverse solution based on test data. The algorithm is iteration-free and superior to the standard pseudo inverse in minimizing the command tracking error.

Pei, Jing↗

NASA NTRS2019

A Finite Volume Procedure for Thermo-Fluid System Analysis in a Flow Network

This paper describes a finite volume procedure for network flow analysis in a thermofluid system. A flow network is defined as a group of inter-connected control volumes called “nodes” that are connected by “branches”. The mass and energy conservation equations are solved at the nodes and momentum conservation equations are solved at the branches. The flow network also includes solid nodes to account for fluid to solid heat transfer. The heat conduction equation is solved at the solid nodes in conjunction with the flow equations. The properties of a real fluid are calculated using a thermodynamic property program and used in the conservation equations. The system of equations describing the fluid-solid network is solved by a hybrid numerical method that is a combination of the Newton-Raphson and successive substitution method. This procedure has been incorporated into a general-purpose computer program, the Generalized Fluid System Simulation Program, GFSSP. This paper also presents the application and verification of the method by comparison with test data for several applications that include (a) internal flow in a rocket engine turbo-pump, (b) pressurization and loading of a cryogenic propellant tank, (c) fluid transient during sudden opening of valve for priming of an evacuated feed line, and (d) chill down of a cryogenic transfer line with phase change and two phase flows.

Majumdar, Alok↗

NASA NTRS2019

Common Habitat for Long Duration Transit and Surface Operations

This concept is based on the Skylab II concept, which proposes using a SLS propellant tank as the primary structure of a habitat. Common Habitat takes this a step further, proposing an interior architecture that is equally viable as a lunar surface hab, transit hab, and Mars surface hab, thus enabling a single development to support all three long duration habitat roles. Common Habitat has been pursued by the PI and co-investigators at MSFC and other centers in a spare time / funding available basis, often with heavy student intern support. Student resources have developed preliminary concepts, but these models must be refined to reflect accurate vehicle subsystems, utilities, stowage, workstations, and other crew systems. This investigation applies NASA civil servant and contractor expertise to correct errors in the most recent student concepts in order to prepare them for future human-in-the-loop evaluations to determine the most viable configuration for a Common Habitat. In prior studies there was never an opportunity to trade both crew size and internal orientation. Prior work has been a series of point designs with inconsistent design constraints, making apples to apples comparisons impossible. The most recent student work produced CAD models of four habitat variants for future use in a trade study, but the student products contain model construction and geometry errors, inaccurate subsystems design and placement, and unrealistic structural elements and outfitting. These must be corrected before the models can be moved into a VR environment for humanin- the-loop testing. Through consultation with SMEs throughout the Agency, the fidelity of these designs will be upgraded to a level of quality sufficient to make a comparative analysis of crew size sensitivity and internal orientation. The product of this activity is four upgraded CAD models, each reflecting a different configuration of the Common Habitat. One is a horizontal configuration using the full length of the SLS LOX tank. One is a vertical configuration using the same tank dimensions. The third is a horizontal configuration using a truncated SLS LOX tank (half the barrel length). The fourth is a vertical configuration of the half-length SLS LOX tank.

Howard, Robert L.↗

NASA NTRS2019

Common Habitat for Long Duration Transit and Surface Operations

Howard, Robert L.↗

NASA NTRS2019

Suppression of Nonlinear Rotary Slosh Dynamics Using the SLS Adaptive Augmenting Control System Demonstration on a Quadcopter Testbed

Liquid propellant makes up a significant portion of the total weight for large launch vehicles such as Saturn V, Space Shuttle, and the Space Launch System. Careful attention must be given to the influence of fuel slosh motion on the stability of the vehicle in the design of the Flight Control System (FCS). Historically, there have been instances where a poorly designed FCS in addition to a lack of passive damping have caused the slosh mass to drive the launch vehicle unstable. The dynamics behind controlling a quadcopter/hanging mass configuration is analogous to that of controlling the attitude of a rocket with a single propellant tank. The quadcopter/hanging mass configuration offers a reasonably accurate platform for assessing the real-time effectiveness of the SLS Adaptive Augmenting Controller in suppressing slosh instability. Flight test experiments were carried out at the NASA Langley Research Center's Autonomy Incubator. During both simulation and flight test, the hanging mass was intentionally made unstable and the adaptive algorithm successfully suppressed the instability as expected.

Pei, Jing↗

NASA NTRSDate not supplied

NASA's Space Launch System Progress Toward the Launch Pad

NASA’s Space Launch System (SLS) took a substantial step toward the launch site in2020 with the move of the Artemis I core stage from the manufacturing site to the test stand as the program plans for a 2021 launch. (Fig. 1) SLS is NASA’s evolvable super-heavy-lift launch vehicle to support deep space exploration. It is based on evolutionary improvement to existing proven propulsion systems. Its twin solid rocket boosters employ a five-segment motor based on the four-segment space shuttle motor. Its four RS-25 main engines will operate at thrust levels higher than those in the space shuttle program. The core stage is anew design that will support propellant tanks for the engines and serve as the attach point for the boosters. Modern streamlined manufacturing processes and materials have been incorporated into the initial configuration with planned onramps for improved performance and/or affordability in subsequent versions. The primary role of SLS is to anchor the transportation for NASA’s Artemis Program to return humans to the Moon and build on the exploration that began during the Apollo Program. This paper will discuss progress to date for the SLS Program and look ahead to important milestones in 2020 and beyond.

John H Honeycutt↗

NASA NTRSDate not supplied

Development and Characterization of the Integrally Stiffened Cylinder (ISC) Process for Launch Vehicles and Aircraft Fuselage Structures

Over the past decade, NASA Langley has led development of the Integrally Stiffened Cylinder (ISC) Process, a near-net-shape, flow-forming technology. This innovative process is being evaluated for launch vehicle and commercial aircraft manufacturing. A thick-walled, cylindrical preform is flow-formed using a single operation into a thin-walled barrel with integral longitudinal stiffeners. The one-piece, stiffened barrels offer a direct replacement for conventional multi-piece, welded or riveted structures. A cost-benefit analysis for launch vehicle cryogenic propellant tanks estimated that the ISC process offers up to a 50% reduction in manufacturing costs and a 10% reduction in mass. NASA, the European Space Agency (ESA), and industry partners have teamed to manufacture 3-m(10-ft.) diameter aluminum ISCs. The intent is to demonstrate the commercial viability of the process at a scale relevant to the commercial launch and aircraft industries. Development activities will be presented, including process scale-up, formability assessment, and mechanical property testing.

ISC, flow forming, near net shape↗

NASA NTRSDate not supplied

Summary of Testing Results for the Structural Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER)

Testing was completed on the Structural Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER) between August 2019 and January 2020. SHIIVER was designed to be a test bed for the scaling of cryogenic fluid management technologies as applied to large upper stages and long duration in-space stages. The baseline SHIIVER design consists of uninsulated structural skirts attached to a propellant tank insulated with polyurethane Spray-On Foam Insulation (SOFI). The initial testing of the SHIIVER hardware was with liquid hydrogen and sought to demonstrate the use of boil-off vapor to intercept heat on a structural skirt, multilayer insulation (MLI) on the tank domes, and the radio frequency mass gauge (RFMG). Testing was completed in four stages: a baseline thermal vacuum test prior to installation of the MLI on the tank domes, a thermal vacuum test after the MLI installation, a reverberant acoustic test, and a subsequent thermal vacuum test to verify that no damage occurred during the reverberant acoustic testing. Each thermal vacuum test with chamber wall at ambient temperature and vacuum level in the 10-6 torr range was conducted continuously between approximately 90% full and 25% full. Test results showed that the vapor cooling reduced the heat load to the tank by approximately 10%, but the boil-off by less than 3% at 50% full, with and without MLI installed on the domes. The MLI installed on the domes reduced the heat load to the tank by approximately 40% at all fill levels, but the boil-off by approximately 25% at 90% full and 45% below 65% full. The RFMG performed well over all fill ranges, and several RF tank modes were used to gauge the mass of fluid in the tank. SHIIVER was then exposed to an acoustic environment of 147 dB OASPL (overall sound pressure level) in a reverberant chamber. The acoustic environment envelopes the upper stage internal acoustic level of several different modern launch vehicles. No structural or thermal performance changes were observed after exposure to the acoustic environment. Final thermal vacuum testing after the acoustic testing showed no degradation to the MLI due to the acoustic environment as measured via system heat loads

SHIIVER↗

NASA NTRS2022

Development of Advanced Manufacturing Approaches for Single-Piece Launch Vehicle Structures

Advanced, near-net shape manufacturing methods have the potential to enable production of structures with fewer welds and reduced machining requirements. Two such methods that are suitable for the manufacturing of single-piece, stiffened barrel-shaped structures are presented. The first solution deployed existing manufacturing technology to produce a thick-walled barrel with integrally machined stiffeners. In this study, an 8-ft. diameter barrel was produced and subjected to a buckling test. Data on the manufacturing and testing of this barrel are provided and compared against the traditional multi-piece weld construction approach. The single-piece barrel resulted in a 28% greater load carrying capacity than the welded barrel. A second solution utilized a novel flow-forming technique to produce the barrel and stiffeners in one process without the need for welding or machining. This is an emerging manufacturing method known as the Integrally Stiffened Cylinder (ISC) process. This innovative process is being evaluated for launch vehicle and commercial aircraft manufacturing. The one-piece, stiffened ISC barrels, which have been successfully fabricated at 10 ft. in diameter, offer a direct replacement for conventional multi-piece, welded or riveted structures. A cost-benefit analysis for launch vehicle cryogenic propellant tanks estimated that the ISC process offers up to a 50% reduction in manufacturing costs and a 10% reduction in mass.

integrally stiffened cylinder (ISC)↗

NASA NTRSDate not supplied

Interfacial thermodynamics of cryogenic fluids: The effect on non-condensable gas on fluid storage

Propellant tanks that contain cryogenic fluids (CFs) in low gravity conditions are usually pressurized with non-condensable (NC) gases for fast extraction. Unfortunately, the presence of NC gases causes CFs to exhibit higher boil-offs compared to pure CF systems. For optimal utilization of the cryogenic fuels, these higher boil-off rates must be minimized. Our goal is to quantify the effects NC gases have on the evaporation and condensation dynamics of CFs to develop strategies that mitigate the higher boil-offs. The hypothesis is that the NC gas accumulates around the CF’s liquid vapor interface in the Knudsen layer, creating a kinetic barrier for both evaporation and condensation. Testing this hypothesis with experimental methods is difficult due to the transient nature of the Knudsen layer, which is only a few nanometers thick. Furthermore, experimental capabilities are limited in low gravity conditions and extremely expensive. Accordingly, we approach this problem from a theoretical perspective. In this study, we employ molecular dynamics (MD) simulations to probe the interfacial mechanisms that affect CF’s evaporation and condensation at varying concentrations of NC gases. Specifically, we use nitrogen (N2) and oxygen (O2) as our CFs and neon (Ne) as our NC gas. Using MD simulations, we show that Ne accumulates at N2 liquid vapor interface across a wide range of Ne concentrations, thereby impeding mass transport of N2. Our simulations allow for direct computation of the molar flux as well as the mass accommodation coefficient (MAC) which can then be used as input parameters to continuum fluid dynamics (CFD) models for optimal storage tank design.

Michael Robert DeLyser↗

NASA NTRSDate not supplied

Evaluating the Performance of Hybrid Vehicles between LEO, Cislunar Space, and Mars

Nuclear power and electric propulsion technologies can enable a broad range of existing and future space mission concepts. This paper evaluates how these two technology classes, when combined into systems with various performance levels (and paired with a chemical propulsion system) affect vehicle sizing and mission feasibility for round-trip, opposition-class, crewed missions to Mars. The technology trade-space evaluated in this paper encompasses: various nuclear power system mass-efficiency levels (or power-specific mass), a range of output power levels available to the electric propulsion system, two electric propulsion thruster technologies (Hall and Magnetoplasmadynamic) at various characteristic specific impulse levels, and a range of propellant tank mass sizing coefficient. Each combination of the technologies is evaluated for the 2039 and the 2042 Mars mission opportunities for a range of imposed total mission duration constraints. The results show that synergies between technology performance levels exist and that increasing the performance of some parameters might have undesired consequences for the trajectory and vehicle. It was found that increasing the mass-efficiency of the nuclear power system is broadly beneficial, resulting in lower vehicle masses and enabling shorter mission durations. However, increasing the specific impulse of a thruster or increasing the available power level do not always yield a net benefit to the vehicle or mission; since increasing the specific impulse or power level (holding power system mass-efficiency constant) tends to decrease the vehicle thrust-to-mass ratio. Furthermore, the results show that mission feasibility is highly dependent on the mission duration; the required technology performance level to enable a mission is relaxed as the mission duration increases. Given the optimally sized vehicles resulting from the trade-space for a crewed mission to Mars, this paper also evaluates the impact of aggregating the Mars mission vehicle in Low-Earth Orbit and transferring it to the Mars mission departure node (using the on-board power and propulsion systems). The trade-space for this portion of the analysis includes the option of in-space vehicle refueling at the Mars mission departure node. The results show that on-orbit refueling can relax the required technology performance levels to achieve lower launch masses and orbit-raising durations at the cost of increased architecture complexity and risk.

Nuclear Electric Propulsion↗

NASA NTRS2023

Development Unit for In-Space Pneumatic Helium Transfer Compressor

High-pressure gaseous helium (GHe) plays a critical role in spacecraft propulsion due to its near exclusive use for pressurization of propellant tanks. Several different designs for helium compressor technologies have been developed for use on Earth; however, significant design changes are required to meet the power, thermal/heat dissipation, vacuum, vibration/shock, size, mass, and efficiency requirements needed for space and launch environments. Currently, no capability exists for mass efficient on-orbit GHe (or xenon) transfer, nor has it ever been attempted. This effort proposed several challenging requirements for the vendor to develop a one-of-a-kind pneumatic compressor prototype. In order to develop a compressor prototype at a TRL 4, a partnership with Air Squared, Inc. was established. This CIF project concluded with the receipt of a prototype along with a proposed design for a flight unit; however, it will not be built due to funding limitations. The vendor had moderate success with initial testing of the prototype at lower pressures (with shimming configured for low pressure). There were delays with testing since the unit was built for higher pressures and the low pressure shimming was challenging to implement. It was found that operation of the compressor at lower pressures caused increased friction on the motor shaft which resulted in the unit drawing more power (current). This was initially thought to be a motor or controller defect. Subsequent high pressure testing at KSC showed that the friction was reduced for the optimal design point (with a high pressure shimming configuration) and the motor (and controller) performed nominally. Following the vendor testing, the unit was tuned/shimmed for a nominal inlet pressure and shipped to KSC for high pressure testing using the custom-designed setup in the Vehicle Assembly Building. The unit performed well, showing the capability to achieve over higher pressure of compression. The pressure differential could likely increase, however, the maximum output of the power supply used for testing was reached, limiting the compressor motor’s capability. Following the GN2 testing, the test setup was converted to GHe with at high pressure. It was shown that although the unit was unable to achieve similar high pressure differentials due to reverse leakage of the GHe through the scroll tip seals, it did achieve the desired GHe compression. This may still be useful for some applications, however, for the purpose of on-orbit GHe transfer, the single stage design would need to be changed to include multiple stages to achieve the compression goals initially proposed by the project.

Brian Nufer↗

NASA NTRS2023

NASA Engineering and Safety Center Technical Bulletin No. 23-02: Safety Considerations when Repurposing Commercially Available Flight Termination Systems from Uncrewed to Crewed Launch Vehicles

Both uncrewed and crewed launch vehicles (LV) require Flight Termination Systems (FTS) for Range Safety to protect the public and ground assets in the event of a LV failure. Flight crew safety in this context is an added consideration for human spaceflight. The FTS is an electroexplosive system that activates destruct charges to rupture propellant tanks and shutdown engines during flight termination. Commercially available FTS units have been developed for uncrewed applications and are now being repurposed to crewed applications. A consequence of using these systems is that they are designed for public and ground crew safety, though inadequate for flight crew safety. Missing are Human Space Flight (HSF) design controls for inadvertent activation during crewed ascent and protection for crew emergency abort.

Flight Termination Systems↗

NASA NTRSDate not supplied

Designing the Imaging and Breadboard System for FBCE-TL Project

This presentation outlines the support work done on designing the imaging and breadboard system of the Transfer Line (TL) project, which is the third module for the Flow Boiling and Condensation Experiment (FBCE) operating in the Fluids Integrated Rack on-board the International Space Station (ISS). The goal of the FBCE is to serve as a facility for flow boiling and condensation investigating two-phase flow and heat transfer data in microgravity. The TL module aims use existing FBCE hardware to study cryogenic chilldown, a large component of cryogenic propellant tank transfer research. The TL test module will perform chilldown studies with normal-perfluorohexane (nPFH), a cryogenic simulant, flowing through a super-heated line, which is heated through the process of Joule-heating. SolidWorks and Creo were used to create electrically conductive pipe junctions that connect and seal several Indium Tin Oxide (ITO) coated sapphire tube sections together to allow for low leak probability fluid flow and circumferentially uniform electrical current density. It is necessary to understand and calculate the thermal expansion of the tubing sections such that the junctions would not leak at anticipated operating temperatures and pressures, would maintain electrical contact, and not crack the crystalline tubes. Optical imaging techniques were used to build an experimentally validated image remapping process for flight-like setups of cylindrical tubes. Then MATLAB was used to create an algorithm that distorts an image based on the refraction of light passing through the mediums we were imaging through to accurately visualize the cryogenic chilldown that is of scientific interest. This research will assist NASA to move forward with long-duration spaceflight missions where knowledge of efficient heat transfer processes becomes more critical.

imaging system↗

NASA NTRSDate not supplied

NASA’s Space Launch System: Artemis I Results and the Path Forward

On Nov. 16, 2022, NASA launched the Artemis I mission on the agency’s new SLS (Space Launch System) rocket. The debut launch of the integrated vehicle sent an uncrewed Orion spacecraft into a distant retrograde (DRO) orbit about the Moon to accomplish multiple mission objectives, including evaluating Orion’s heatshield at lunar return velocities and collecting critical data from the first flight of the new superheavy lift launch vehicle. Post-flight data analyses show that the rocket performed with a high degree of precision and accuracy. In addition to completing testing and launch of the Artemis I SLS in 2022, notable progress was made on hardware for future launch vehicles that will support Artemis missions II, III, IV, and V. The Artemis II core stage is mostly complete, and its RS-25 engines were integrated in the fall of 2023. The Artemis II solid rocket booster motor segments were shipped to the launch site in September 2023. Propellant tanks are being manufactured for additional missions, solid booster segments are being cast and prepared, software is being developed, and production of new RS-25 engines is underway. A more capable upper stage is in development, as well as a large adapter to encapsulate a 10-metric ton (t) co-manifested payload. This paper will provide details on the Artemis I flight as well as cover hardware progress for the future flights as available.

John Honeycutt↗

NASA NTRSDate not supplied

Validation of Thermodynamic Behavior of Liquid Propellants under Sloshing and Draining

In recent years, considerable effort has been devoted to studying the future use of liquid methane (LCH4) in land, air, and space vehicle applications because of its high density and handling characteristics. This work presents a validation of computational tools for the thermodynamics characterization of a propellant tank undergoing sloshing and draining-induced thermal destratification as part of our continuous effort to improve simulation capabilities for support of NASA’s current and future flight programs. A multi-phase computational fluid dynamics (CFD) code developed at NASA MSFC, Loci/STREAM-VOF, is applied to predict gaseous pressurant requirements during the ramping, holding, and draining phases of operation with liquid methane. The experimental work conducted at the NASA K-site facility is used for validation. The effort showed that Loci/STREAM-VOF is capable of capturing the important findings from the previous experiments: (1) the pressurant mass required increases with the expulsion time due to longer mass transfer time at the interface; and (2) the pressurant mass required decreases with the increase in inlet temperature. Comparison with experimental data shows consistent good agreement at different expulsion times and different inlet gas temperatures.

H. Q. Yang↗

NASA NTRSDate not supplied

Green Run Modal Test of the NASA Space Launch System Core Stage

The Core Stage of the new NASA Space Launch System (SLS) is a 212-foot-tall rocket assembly—consisting primarily of two cryogenic propellant tanks, an engine section, and four RS-25 rocket engines—that will send crew and large payloads to the moon and beyond for NASA’s Artemis program. Prior to SLS assembly, the Core Stage completed a series of structural and functional tests in the B-2 Test Stand at Stennis Space Center, designated Green Run. The goal of Green Run was to verify analytical models, confirm proper subsystem operation, and test-fire all four RS-25 engines of the Core Stage. In January 2020, Green Run testing began with an experimental modal analysis test, performed by the Marshal Space Flight Center modal test team. A free-boundary test condition of the Core Stage was simulated as close as possible by suspending the massive launch vehicle from the B-2 Test Stand crane. Modal excitation was provided by a pair of 250-lb electro-dynamic shakers for multi-shaker random vibration testing, as well as a 12-pound instrumented hammer for impact testing. Modal response was measured with 550 accelerometer channels distributed on both the Core Stage and the B-2 Test Stand derrick crane. Following one very long day of testing, frequency response functions were calculated from the measured time histories in the target mode frequency range of 5 Hz to 15 Hz, and mode shapes, frequencies, and damping values were successfully estimated. The case-study presented in this paper will discuss Green Run, the SLS Core Stage, the modal test setup and execution, as well as a brief overview of the test results. Challenges associated with testing such a large, suspended structure in an outdoor environment will be discussed as well.

Space Launch System↗

NASA NTRSDate not supplied

Development of Advanced Lightweight Polymeric Foam Thermal Protection for Space Launch Vehicles

Lightweight thermal protective materials are essential for temperature regulation during ascent of space launch vehicles, especially for expendable launch vehicles and cryogenic propellant tanks. Among the lightest thermal protective materials commonly in use is polymeric foam, in the form of pour foam or spray-on foam insulation (SOFI). Foam processing is rapid, can accommodate large-scale or small-scale challenging geometries easily, and can be used in a climate controlled or non-climate-controlled setting. Application systems can be fixed or portable. Use-temperatures of foam range from cryogenic to those that will ablate the material.

TPS↗