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NASA NTRS2013

Proposed Flight Research of a Dual-Bell Rocket Nozzle Using the NASA F-15 Airplane

For more than a half-century, several types of altitude-compensating rocket nozzles have been proposed and analyzed, but very few have been adequately tested in a relevant flight environment. One type of altitude-compensating nozzle is the dual-bell rocket nozzle, which was first introduced into literature in 1949. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. This paper proposes a method for conducting testing and research with a dual-bell rocket nozzle in a flight environment. We propose to leverage the existing NASA F-15 airplane and Propulsion Flight Test Fixture as the flight testbed, with the dual-bell nozzle operating during captive-carried flights, and with the nozzle subjected to a local flow field similar to that of a launch vehicle. The primary objective of this effort is not only to advance the technology readiness level of the dual-bell nozzle, but also to gain a greater understanding of the nozzle mode transitional sensitivity to local flow-field effects, and to quantify the performance benefits with this technology. The predicted performance benefits are significant, and may result in reducing the cost of delivering payloads to low-Earth orbit.

Jones, Daniel S.↗

NASA NTRS2013

Proposed Flight Research of a Dual-Bell Rocket Nozzle Using the NASA F-15 Airplane

Jones, Daniel S.↗

NASA NTRS2017

A CubeSat-Payload Radiation-Reliability Assurance Case using Goal Structuring Notation

CubeSats have become an attractive platform for universities, industry, and government space missions because they are cheaper and quicker to develop than full-scale satellites. One way CubeSats keep costs low is by using commercial off-the-shelf parts (COTS) instead of space-qualified parts. Space-qualified parts are often costlier, larger, and consume more power than their commercial counterparts precluding their use within the CubeSat form-factor. Given typical power budgets, monetary budgets, and timelines for CubeSat missions, conventional radiation hardness assurance, like the use of space-qualified parts and radiation testing campaigns of COTS parts, is not practical. Instead, a system-level approach to radiation effects mitigation is needed. In this paper an assurance case for a system-level approach to mitigate radiation effects of a CubeSat science experiment is expressed using Goal Structuring Notation (GSN), a graphical argument standard. The case specifically looks at three main mitigation strategies for the radiation environment: total ionizing dose (TID) screening of parts, detection and recovery from single-event latch-ups (SEL) and single-event functional interrupts (SEFI). The graphical assurance case presented makes a qualitative argument for the radiation reliability of the CubeSat experiment using part and system-level mitigation strategies.

COTS↗

NASA NTRS1981

A study to define the research and technology requirements for advanced turbo/propfan transport aircraft

The feasibility of the propfan relative to the turbofan is summarized, using the Douglas DC-9 Super 80 (DS-8000) as the actual operational base aircraft. The 155 passenger economy class aircraft (31,775 lb 14,413 kg payload), cruise Mach at 0.80 at 31,000 ft (8,450 m) initial altitude, and an operational capability in 1985 was considered. Three propfan arrangements, wing mounted, conventional horizontal tail aft mounted, and aft fuselage pylon mounted are selected for comparison with the DC-9 Super 80 P&WA JT8D-209 turbofan powered aircraft. The configuration feasibility, aerodynamics, propulsion, structural loads, structural dynamics, sonic fatigue, acoustics, weight maintainability, performance, rough order of magnitude economics, and airline coordination are examined. The effects of alternate cruise Mach number, mission stage lengths, and propfan design characteristics are considered. Recommendations for further study, ground testing, and flight testing are included.

Goldsmith, I. M.↗

NASA NTRS1992

The 10 kW power electronics for hydrogen arcjets

A combination of emerging mission considerations such as 'launch on schedule', resource limitations, and the development of higher power spacecraft busses has resulted in renewed interest in high power hydrogen arcjet systems with specific impulses greater than 1000 s for Earth-space orbit transfer and maneuver applications. Solar electric propulsion systems with about 10 kW of power appear to offer payload benefits at acceptable trip times. This work outlines the design and development of 10 kW hydrogen arcjet power electronics and results of arcjet integration testing. The power electronics incorporated a full bridge switching topology similar to that employed in state of the art 5 kW power electronics, and the output filter included an output current averaging inductor with an integral pulse generation winding for arcjet ignition. Phase shifted, pulse width modulation with current mode control was used to regulate the current delivered to arcjet, and a low inductance power stage minimized switching transients. Hybrid power Metal Oxide Semiconductor Field Effect Transistors were used to minimize conduction losses. Switching losses were minimized using a fast response, optically isolated, totem-pole gate drive circuit. The input bus voltage for the unit was 150 V, with a maximum output voltage of 225 V. The switching frequency of 20 kHz was a compromise between mass savings and higher efficiency. Power conversion efficiencies in excess of 0.94 were demonstrated, along with steady state load current regulation of 1 percent. The power electronics were successfully integrated with a 10 kW laboratory hydrogen arcjet, and reliable, nondestructive starts and transitions to steady state operation were demonstrated. The estimated specific mass for a flight packaged unit was 2 kg/kW.

Hamley, John A.↗

NASA NTRS2012

Phase Function Determination in Support of Orbital Debris Size Estimation

To recover the size of a space debris object from photometric measurements, it is necessary to determine its albedo and basic shape: if the albedo is known, the reflective area can be calculated; and if the shape is known, the shape and area taken together can be used to estimate a characteristic dimension. Albedo is typically determined by inferring the object s material type from filter photometry or spectroscopy and is not the subject of the present study. Object shape, on the other hand, can be revealed from a time-history of the object s brightness response. The most data-rich presentation is a continuous light-curve that records the object s brightness for an entire sensor pass, which could last for tens of minutes to several hours: from this one can see both short-term periodic behavior as well as brightness variations with phase angle. Light-curve interpretation, however, is more art than science and does not lend itself easily to automation; and the collection method, which requires single-object telescope dedication for long periods of time, is not well suited to debris survey conditions. So one is led to investigate how easily an object s brightness phase function, which can be constructed from the more survey-friendly point photometry, can be used to recover object shape. Such a recovery is usually attempted by comparing a phase-function curve constructed from an object s empirical brightness measurements to analytically-derived curves for basic shapes or shape combinations. There are two ways to accomplish this: a simple averaged brightness-versus phase curve assembled from the empirical data, or a more elaborate approach in which one is essentially calculating a brightness PDF for each phase angle bin (a technique explored in unpublished AFRL/RV research and in Ojakangas 2011); in each case the empirical curve is compared to analytical results for shapes of interest. The latter technique promises more discrimination power but requires more data; the former can be assembled in its essentials from fewer measurements but will be less definitive in its assignments. The goal of the present study is to evaluate both techniques under debris survey conditions to determine their relative performance and, additionally, to learn precisely how a survey should be conducted in order to maximize their performance. Because the distendedness of objects has more of an effect than their precise shape in calculating a characteristic dimension, one is interested in the techniques discrimination ability to distinguish between an elongated rectangular prism and a short rectangular prism or cube, or an elongated cylinder from a squat cylinder or sphere. Sensitivity studies using simulated data will be conducted to determine discrimination power for both techniques as a function of amount of data collected and range (and specific region) of phase angles sampled. Empirical GEODSS photometry data for distended objects (dead payloads with solar panels, rocket bodies) and compact objects (cubesats, calibration spheres, squat payloads) will also be used to test this discrimination ability. The result will be a recommended technique and data collection paradigm for debris surveys in order to maximize this type of discrimination.

Hejduk, M. D.↗

NASA NTRS2011

Z-Pinch Magneto-Inertial Fusion Propulsion Engine Design Concept

Fusion-based nuclear propulsion has the potential to enable fast interplanetary transportation. Due to the great distances between the planets of our solar system and the harmful radiation environment of interplanetary space, high specific impulse (Isp) propulsion in vehicles with high payload mass fractions must be developed to provide practical and safe vehicles for human spaceflight missions. Magneto-Inertial Fusion (MIF) is an approach which has been shown to potentially lead to a low cost, small fusion reactor/engine assembly (1). The Z-Pinch dense plasma focus method is an MIF concept in which a column of gas is compressed to thermonuclear conditions by an estimated axial current of approximately 100 MA. Recent advancements in experiments and the theoretical understanding of this concept suggest favorable scaling of fusion power output yield as I(sup 4) (2). The magnetic field resulting from the large current compresses the plasma to fusion conditions, and this is repeated over short timescales (10(exp -6) sec). This plasma formation is widely used in the field of Nuclear Weapons Effects (NWE) testing in the defense industry, as well as in fusion energy research. There is a wealth of literature characterizing Z-Pinch physics and existing models (3-5). In order to be useful in engineering analysis, a simplified Z-Pinch fusion thermodynamic model was developed to determine the quantity of plasma, plasma temperature, rate of expansion, energy production, etc. to calculate the parameters that characterize a propulsion system. The amount of nuclear fuel per pulse, mixture ratio of the D-T and nozzle liner propellant, and assumptions about the efficiency of the engine, enabled the sizing of the propulsion system and resulted in an estimate of the thrust and Isp of a Z-Pinch fusion propulsion system for the concept vehicle. MIF requires a magnetic nozzle to contain and direct the nuclear pulses, as well as a robust structure and radiation shielding. The structure, configuration, and materials of the nozzle must meet many severe requirements. The configuration would focus, in a conical manner, the Deuterium-Tritium (D-T) fuel and Lithium-6/7 liner fluid to meet at a specific point that acts as a cathode so the Li-6 can serve as a current return path to complete the circuit. In addition to serving as a current return path, the Li liner also serves as a radiation shield. The advantage to this configuration is the reaction between neutrons and Li-6 results in the production of additional Tritium, thus adding further fuel to the fusion reaction and boosting the energy output. To understand the applicability of Z-Pinch propulsion to interplanetary travel, it is necessary to design a concept vehicle that uses it. The propulsion system significantly impacts the design of the electrical, thermal control, avionics, radiation shielding, and structural subsystems of a vehicle. The design reference mission is the transport of crew and cargo to Mars and back, with the intention that the vehicle be reused for other missions. Several aspects of this vehicle are based on a previous crewed fusion vehicle study called Human Outer Planet Exploration (HOPE), which employed a Magnetized Target Fusion (MTF) propulsion concept. Analysis of this propulsion system concludes that a 40-fold increase of Isp over chemical propulsion is predicted. This along with a greater than 30% predicted payload mass fraction certainly warrants further development of enabling technologies. The vehicle is designed for multiple interplanetary missions and conceivably may be suited for an automated one-way interstellar voyage.

Miernik, Janie H.↗

NASA NTRS1984

Space reliability technology - A historical perspective

The progressive improvements in reliability of launch vehicles is traced from the Vanguard rocket to the STS. The Vanguard, built with minimal redundancy and a high mass ratio, was used as an operational vehicle midway through its test program in an attempt to meet the perceived challenge represented by the Sputnik. The fourth Vanguard failed due to inadequate contamination prevention and lack of inspection ports. Automatic firing sequences were adopted for the Titan rockets, which were an order of magnitude larger than the Vanguard and therefore had room for interior inspections. Qualification testing and reporting were introduced for components, along with X ray inspection of fuel tank welds. Dual systems were added for flight critical components when the Titan became man-rated for the Gemini program. Designs incorporated full failure mode effects and criticality analyses for the Apollo program, which exposed the limits of applicability of numerical reliability models. Fault tree analyses and program milestone reviews were initiated. The worth of man-in-the-loop in space activities for reliability was demonstrated with the rescue of Skylab after solar panel and meteoroid shield failures. It is now the reliability of the payload, rather than the vehicle, that is questioned for Shuttle launches.

Cohen, H.↗

NASA NTRS2011

Legacy of Biomedical Research During the Space Shuttle Program

The Space Shuttle Program provided many opportunities to study the role of spaceflight on human life for over 30 years and represented the longest and largest US human spaceflight program. Outcomes of the research were understanding the effect of spaceflight on human physiology and performance, countermeasures, operational protocols, and hardware. The Shuttle flights were relatively short, < 16 days and routinely had 4 to 6 crewmembers for a total of 135 flights. Biomedical research was conducted on the Space Shuttle using various vehicle resources. Specially constructed pressurized laboratories called Spacelab and SPACEHAB housed many laboratory instruments to accomplish experiments in the Shuttle s large payload bay. In addition to these laboratory flights, nearly every mission had dedicated human life science research experiments conducted in the Shuttle middeck. Most Shuttle astronauts participated in some life sciences research experiments either as test subjects or test operators. While middeck experiments resulted in a low sample per mission compared to many Earth-based studies, this participation allowed investigators to have repetition of tests over the years on successive Shuttle flights. In addition, as a prelude to the International Space Station (ISS), NASA used the Space Shuttle as a platform for assessing future ISS hardware systems and procedures. The purpose of this panel is to provide an understanding of science integration activities required to implement Shuttle research, review biomedical research, characterize countermeasures developed for Shuttle and ISS as well as discuss lessons learned that may support commercial crew endeavors. Panel topics include research integration, cardiovascular physiology, neurosciences, skeletal muscle, and exercise physiology. Learning Objective: The panel provides an overview from the Space Shuttle Program regarding research integration, scientific results, lessons learned from biomedical research and countermeasure development.

Hayes, Judith C.↗

NASA NTRSDate not supplied

22 N HPGP Thruster Life Testing

In the ever-changing paradigm of efficient and capable spacecraft design, scientific missions continue pushing the envelope enabling spacecraft subsystems to deliver effective solutions to meet challenging new mission/spacecraft needs. From an in-space storable liquid chemical propulsion perspective, monopropellant hydrazine has been, and continues to be, a dependable propellant with considerable flight heritage, a variety of engine thrust classes available from multiple vendors, with repeatable and reliable performance. Additionally, the space propulsion industry has learned to successfully handle hydrazine, its regulations, the safety protocols, the personnel protective equipment, and the unique training standards–all requisite for loading spacecraft propulsion systems with toxic hypergolic hydrazine. The question now arises as to “what is next for in-space chemical propulsion?” Further, with the evolution and concrete advancements in innovative in-space green propellant technologies, capable of providing realizable benefits to scientific missions, concern over the reliability and availability of this higher performing and safer to handle class of propellants is waning. As science missions move forward with the potential flight in fusion of High Performance Green Propulsion (HPGP), NASA and its industry partners are working to address any gaps in system reliability, performance, or unique operational considerations. Propellant technology that offers both higher performance and significant reduction in personnel hazards compared to hydrazine presents an attractive propulsion subsystem design opportunity. Increased propulsion subsystem performance can result in lower spacecraft launch mass, larger scientific payloads, or extended on-orbit lifetimes. Mission trades using green propulsion technologies have been documented on multiple NASA Goddard Space Flight Center (GSFC) mission classes, examining various parameters and requirements to support mission architectures in Low Earth Orbit (LEO), High Earth Orbit (HEO), geostationary, lunar, planetary, and Quasi-halo orbit around Sun-Earth Lagrange point (L2). The results of these trade studies show promising, attainable benefits. The perceived programmatic risk of flying a newer propulsion technology has, unfortunately, not outweighed the benefits to date. To take advantage of the improved performance and mitigate programmatic risk, HPGP engines must demonstrate life testing at higher propellant throughputs than have currently been demonstrated. In an effort to proactively address the challenges with technology infusion into a risk-averse community, NASA and the Swedish National Space Agency (SNSA) outlined a collaborative Implementing Arrangement (IA) for the respective agencies to pursue increased HPGP technology maturation. This initial IA effort began in 2013, fresh off the heels of the successful PRISMA HPGP technology demonstration mission. The IA targeted objective is to reduce risk to potential future HPGP missions and fully characterize the LMP-103S propellant and associated engine performance. Over the past eight years, HPGP has flown in propulsion systems on twenty-five(25) spacecraft from seven(7) different Launch Ranges around the globe and on seven (7) different Launch Vehicles. Six(6) of these launches involved multiple loading operations for multiple spacecraft. For U.S. Range operations, nine (9) HPGP systems have been processed at Vandenberg Space Force Base(VSFB):six(6) in 2017, and three (3) in 2018. Six (6) more have been processed at Cape Canaveral Air Force Station (CCAFS)in May 2020, with three (3) systems launched in June 2020 and the remaining three (3) system were left loaded and ready until their launch in August of 2020. Three (3) more systems have been processed at Wallops Flight Facility(WFF)and launched in June 2021. In addition, these propulsion subsystems employed heritage propulsion subsystem component such as valves, filters, and pressure transducers, and have further demonstrated nominal functionality in both diaphragm and Propellant Management Device (PDM) propellant tanks. Based on these successes, HPGP technology continues to be considered for NASA Science Mission Directorate missions at GSFC. The work presented herein represents many years of development and collaborative efforts to successfully align higher performance, low toxicity hydrazine alternatives into scientific missions. NASA GSFC Propulsion Engineering, in collaboration with Bradford ECAPS, has developed mission specific thruster design and testing requirements to establish GSFC’s desired test conditions and firing sequences.In2017, the first flight-like 22N HPGP thruster Engineering Qualification Model (EQM-1)was designed and built by Bradford ECAPS to prove out the thruster design, materials, build process, and test campaign with respect to NASA GSFC critical component and mission requirements. This test program was developed to comprehensively test the thruster, the technology, and ultimately increase the 22N HPGP Technology Readiness Level(TRL). EQM-1was tested to environmental qualification levels prior to hot fire performance testing to represent the relevant end-to-end environment (launch to on-orbit operation)with required margin. This thruster demonstrated steady-state and pulse mode operational capability with propellant thruster throughput up to~53kg.At this throughput level, the EQM-1 engine began to present off-nominal performance and the test campaign was halted to allow for non-destructive testing and identify the root cause for the an omalous performance. Capitalizing on the successful elements of the EQM-1 campaign, an upgraded 22N HPGP EQM-2 has been manufactured by Bradford ECAPS to meet the complete GSFC requirements. The EQM-2 thruster’s test campaign has further demonstrated the robustness of the HPGP propulsion technology and increased the Technology Readiness Level (TRL) by undergoing a full acceptance test program, then proceeding into qualification, including environmental testing (vibration and shock to qualification levels),as well as hot-fire life testing, operating at steady-state and pulse modes with increased propellant thruster throughput to~150kg. The HPGP thruster performance testing enables effective HPGP thruster readiness evaluation to meet NASA candidate mission requirements in the future.

High↗

NASA NTRS2003

Deformation and Flexibility Equations for Idealized ARIS Umbilicals, Under Planar End-Loading Conditions

The International Space Station (ISS) relies on the Active Rack Isolation System (ARIS) as the central component of an integrated, station-wide strategy to isolate microgravity space-science experiments. ARIS uses electromechanical actuators to isolate an International Standard Payload Rack (ISPR) from disturbances due to the motion of the ISS. Disturbances to microgravity experiments on ARIS-isolated racks are primarily transmitted via the ARTS power and vacuum umbilicals. Recent experimental tests indicate that these umbilicals resonate at frequencies outside the ARIS controller's bandwidth, at levels of potential concern for certain microgravity experiments. Reduction in the umbilical resonant frequencies could help to address this issue. This report develops equations for the in-plane deflections and flexibilities of an idealized umbilical (thin, flexible, cantilever beam) under end-point, in-plane loading (inclined-force and moment). The effect of gravity is neglected due to the on-orbit application. The analysis assumes an initially straight, cantilevered umbilical with uniform cross-section, which undergoes large deflections with no plastic deformation, such that the umbilical terminus remains in a single quadrant and the umbilical slope changes monotonically. The analysis is applicable to the ARIS power and vacuum umbilicals, under the indicated assumptions.

Hampton, R. David↗

NASA NTRS2000

Deformation and Flexibility Equations for Idealized ARIS Umbilicals, Under Planar End-Loading Conditions

The International Space Station (ISS) relies on the Active Rack Isolation System (ARIS) as the central component of an integrated, station-wide strategy to isolate microgravity space-science experiments. ARIS uses electromechanical actuators to isolate an International Standard Payload Rack (ISPR) from disturbances due to the motion of the ISS. Disturbances to microgravity experiments on ARIS-isolated racks are primarily transmitted via the ARIS power and vacuum umbilicals. Recent experimental tests indicate that these umbilicals resonate at frequencies outside the ARIS controller's bandwidth. at levels of potential concern for certain microgravity experiments. Reduction in the umbilical resonant frequencies could help to address this issue. This paper develops equations for the in-plane deflections and flexibilities of an idealized umbilical (thin, flexible, cantilever beam) under end-point, in-plane loading (inclined-force and moment). The effect of gravity is neglected due to the on:orbit application. The analysis assumes an initially straight. cantilevered umbilical with uniform cross-section. which undergoes large deflections with no plastic deformation, such that the umbilical terminus remains in a single quadrant and the umbilical slope changes monotonically. The analysis is applicable to the ARIS power and vacuum umbilicals. under the indicated assumptions.

Hampton, R. David↗

NASA NTRS1991

Dynamic response of a hammerhead launch vehicle wind-tunnel model

NASA-Langley has wind tunnel-tested an aeroelastically-scaled (1/10th-scale) Atlas-Centaur I Large Payload Fairing launch-vehicle model capable of simulating either of the first two bending vibration modes of the full-scale vehicle, on the basis of a partial-mode technique. While the primary emphasis was on the vehicle's buffet response, angle-of-attack studies were conducted for several payload fairing configurations with a view to both the buffet response and the dynamic stability of off-design conditions. No dynamic instabilities were discovered among the range of configurations, and payload fairing L/D variations were found to have only small buffet effects except for the smallest such value, in the second bending mode configuration.

Cole, Stanley R.↗

NASA NTRS1998

Forced Flow Flame Spreading Test: Preliminary Findings From the USMP-3 Shuttle Mission

The Forced Flow Flame spreading Test (FFFT) is a study of flame spreading over solid fuels in very low-speed air flows. The FFFT experiment is part of research entitled Solid Inflammability Boundary at Low Speeds, (SIBAL) intended for operations on the Space Station. In the FFFT experiment, a series of 15 experiments conducted aboard the space shuttle during the United States Microgravity Payload (USMP-3) mission provided information about the structure and spreading characteristics of flames in low-speed, concurrent flows. The test samples included flat sheets of cellulose and cast cylinders of cellulose, burned in air at velocities of approximately 1 to 8 cm/sec. The test results have been successfully compared to theoretical predictions of the SIBAL program, a fundamentally based numerical simulation of concurrent flow flame spread. Additionally, some guidance for the design characteristics of the SIBAL flight experiment have been obtained including some verification of the theoretical predictions of flame size versus the required size of the SIBAL flow duct, and the effect of the presence of thermocouples in the vicinity of near-limit flames in microgravity.

Sacksteder, Kurt R.↗

NASA NTRS1999

CVD Rhenium Engines for Solar-Thermal Propulsion Systems

Solar-thermal upper-stage propulsion systems have the potential to provide specific impulse approaching 900 seconds, with 760 seconds already demonstrated in ground testing. Such performance levels offer a 100% increase in payload capability compared to state-of-the-art chemical upper-stage systems, at lower cost. Although alternatives such as electric propulsion offer even greater performance, the 6- to 18- month orbital transfer time is a far greater deviation from the state of the art than the one to two months required for solar propulsion. Rhenium metal is the only material that is capable of withstanding the predicted thermal, mechanical, and chemical environment of a solar-thermal propulsion device. Chemical vapor deposition (CVD) is the most well-established and cost-effective process for the fabrication of complex rhenium structures. CVD rhenium engines have been successfully constructed for the Air Force ISUS program (bimodal thrust/electricity) and the NASA Shooting Star program (thrust only), as well as under an Air Force SBIR project (thrust only). The bimodal engine represents a more long-term and versatile approach to solar-thermal propulsion, while the thrust-only engines provide a potentially lower weight/lower cost and more near-term replacement for current upper-stage propulsion systems.

Williams, Brian E.↗

NASA NTRSDate not supplied

Launch Vehicle and Spacecraft Transmitter Electromagnetic Compatibility

The Launch Services Program at NASA’s Kennedy Space Center (KSC) is the primary gate for acquiring commercial vehicles to provide a cost effective ride to space for NASA spacecraft. With the lunar Gateway, more human tended elements are planned for launch. One challenge facing the space industry is the proliferation of communication and science transmitters at frequencies beyond the qualification of space avionics and instruments. Studies have ensued to examine the intricacies of performing radiated susceptibility testing and analysis above 18 GHz. Changes in the launch vehicle communications interface to the range have also lead to new launch vehicle antenna systems and more reliance on GPS and telemetry systems. Finally, research initiated at KSC in the area of predicted electric field distributions in launch vehicle payload fairings have spawned Small Business Technology Transfer initiatives for industry to investigate statistical algorithm and computational improvements in large payload fairing modeling of transmitters at frequencies in the GHz range. These topics, along with electromagnetic compatibility testing for launch vehicles will be discussed.

Electromagnetic Testing Cavity Statistics↗

NASA NTRS2024

On the Effectiveness of Using Woven Outer Plies to Suppress Hole Drilling Face Sheet Damage in Honeycomb Sandwich Structure

A payload adaptor fitting (PAF) structure has been manufactured at NASA’s Marshall Space Flight Center (MSFC). The PAF structure is essentially a large cone made of carbon/epoxy sandwich structure. The PAF structure contains numerous (about 1000) holes at the top and bottom to secure the carbon/epoxy structure in aluminum rings with bolts. In an effort to obtain higher quality holes, both sides of the PAF structure are covered with an additional, non-structural, ply of woven carbon/epoxy placed at ±45⁰ to the vertical direction. This study evaluated the quality of the holes on sandwich structure both with and without the outer woven carbon/epoxy plies via open hole compression (OHC) strength testing. Results showed a 14% increase in strength of the cloth covered sandwich structure, which did not surpass the 18% added weight due to the outer woven plies.

sacrificial plies↗

NASA NTRS1995

The VOrtex Ring Transit EXperiment (VORTEX) GAS project

Get Away Special (GAS) payload G-093, also called VORTEX (VOrtex Ring Transit EXperiment), is an investigation of the propagation of a vortex ring through a liquid-gas interface in microgravity. This process results in the formation of one or more liquid droplets similar to earth based liquid atomization systems. In the absence of gravity, surface tension effects dominate the drop formation process. The Shuttle's microgravity environment allows the study of the same fluid atomization processes as using a larger drop size than is possible on Earth. This enables detailed experimental studies of the complex flow processes encountered in liquid atomization systems. With VORTEX, deformations in both the vortex ring and the fluid surface will be measured closely for the first time in a parameters range that accurately resembles liquid atomization. The experimental apparatus will record images of the interactions for analysis after the payload has been returned to earth. The current design of the VORTEX payload consists of a fluid test cell with a vortex ring generator, digital imaging system, laser illumination system, computer based controller, batteries for payload power, and an array of housekeeping and payload monitoring sensors. It is a self-contained experiment and will be flown on board the Space Shuttle in a 5 cubic feet GAS canister. The VORTEX Project is entirely run by students at the University of Michigan but is overseen by a faculty advisor acting as the payload customer and the contact person with NASA. This paper summarizes both the technical and programmatic aspects of the VORTEX Project.

Bilen, Sven G.↗