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To sustain the extra-vehicular activity (EVA) rate required to assemble and maintain the International Space Station (ISS), we must enhance our ability to plan, train for, and execute EVAs. An underlying analysis capability must be in place to ensure EVA access to all external worksites either as a starting point for ground training, to generate information needed for on-orbit training, or to react quickly to develop contingency EVA plans, techniques, and procedures. This paper describes a potential flight experiment for application of custom human modeling analysis to plan and train for EVAs to enhance space station functionality and usability through assembly and operation.

This paper describes the design and testing of a three phase, 200 Amp. per phase, AC power controller intended to replace electromechanical bus tie and cross tie contactors in commercial aircraft electric power systems. In order to design an effective overall electric power system, both the primary transmission subsystem and the secondary distribution subsystem must operate together, controlling the flow of power in a seamless fashion. This is not possible using electromechanical contactors in the primary subsystem.

The McDonnell Douglas Delta family of launch vehicles, in its more than 30-year history, has proven to be the most reliable spacecraft deployment platform for both the US government and the private sector. This success is due to the continuous and focused application of advanced, affordable engineering and manufacturing technologies in all stages of the design, fabrication, assembly, quality assurance, and launch. One of the recent technological breakthroughs that has enhanced the Delta's service capabilities is the development and use of large composite structures in critical components. Among these structures is the payload fairing, which acts as a protective shroud for the spacecraft. Traditional composite manufacturing techniques, however, are very labor-intensive and time-consuming.

The KEEP EAGLE flight test program was conducted from August 1994 until August 1995 at Edwards AFB by a combined government/contractor test team to evaluate improvements to F-15E high angle-of-attack and spin recovery characteristics. This paper will trace the program from its inception in 1992 until conclusion in 1995, with emphasis on the test approach and flight test techniques employed for this high risk program. Specifically, the test approach included novel assessments of spin recovery control power early in the flight test program using controlled build-ups in yaw rate. The program also used simulation effectively to improve test efficiency and maintain test team proficiency with normal and emergency procedures. These techniques allowed a relatively aggressive flight test program without compromising safety. A total of 18 different aircraft configurations were successfully tested, with 146 developed spins completed throughout the course of 81 program flights.

This paper is focusing on considerations pertaining to riveting/fastening systems and assembly methodology currently in use for large aircraft fuselage structures. Discussion of process principles on which current systems are based is addressing distribution of rivets along the aircraft structure, riveting/fastening systems and equipment flexibility. An attempt was made to predict the most probable future equipment development trends based on the need for more efficiency in all aircraft structural assembly and in high level and final assembly areas.

A detailed SINDA-FLUINT thermo-hydraulic math model of the International Space Station Alpha single-phase thermal control system has been developed to evaluate the system steady state and transient responses. The model is being used to determine critical system performance characteristics, such as line sizing, flow distribution, and temperatures of critical components. It is also used to support the design of the control system required to maintain set point temperature and a constant system pressure. In the future the model will be correlated with test data to provide a reliable tool to support the Space Station operation. A detailed description of the model is presented in this paper, together with sample calculations representing critical Space Station operating conditions. Sensitivity of computed results to variations in critical design parameters is also presented. Since the model will continue to evolve and be improved, logical process to be followed is also outlined.

Tactical aircraft have tire lives as low as 3-5 landings per tire causing excessive support costs. The goal of the Improved Tire Life (ITL) program was to begin developing technology to double aircraft tire life, particularly for tactical aircraft. ITL examined not only the tire, but also aircraft/landing gear design, aircraft operations, and the operational environment. ITL had three main thrusts which were successfully accomplished: 1) development of an analytical tire wear model, 2) initiation of technology development to increase tire life, and 3) exploration of new and unique testing methods for tire wear. This paper reports the work performed and the results of the USAF sponsored ITL program.

This paper describes the capabilities of using a time domain analysis to simulate the electrical waveforms on the MIL-STD-1553B data bus using the circuit simulation program SPICE. This simulation was developed to analyze the various data bus architectures of the Space Station Freedom (SSF) propulsion module system. The advantages of this model over frequency domain models is that the waveform is directly calculated, not synthesized, providing a more accurate and detailed waveform representation. Also, nonstandard configurations and effects can be modeled with accurate simulation results. The output of the simulation is an electrical waveform which can be measured at any point in the bus architecture. Therefore it is possible to see the distortion effects on the transmitted signal's waveform at any point on the MIL-STD-1553B data bus. This model accurately simulates the stray impedance effects of all bus components, and is flexible enough to examine any transmitted waveform.

A CELSS has been defined based on current or near-term technology. The CELSS was sized to support the metabolic load of four people on the Moon for ten years. A metabolic load of 14 MJ/person/day is assumed, including an average of 2.6 hr of EVA/person/day. Close to 100% closure of water, and oxygen, and 85% closure of the food loop is assumed. With 15% of the calories supplied from Earth, this should provide adequate dietary variety for the crew along with vitamin and mineral requirements. Other supply and waste removal requirements are addressed. The basic shell used is a Space Station Freedom 7.3 m (24 ft) module. This is assumed to be buried in regolith to provide protection from radiation, meteoroids, and thermal extremes. A solar dynamic power system is assumed, with a design life of 10 years delivering power at 368 kWh/kg. Initial estimates of size are that 73 m2 of plant growth area are required, giving a plant growth volume of about 73 m3.

A method for applying an organic coating to Z-93, an inorganic white thermal control paint, was developed to protect Z-93 from contamination and damage. A layer of FEP Teflon™ was applied over Z-93 to provide a smooth, continuous surface without adversely affecting its optical properties. Additionally, new low-absorptance, controlled-emittance thermal control paints were developed for low Earth orbit (LEO) applications, such as the International Space Station. These paints have a range of infrared emittances from 0.26 to 0.88, and are stable in simulated LEO environments, including atomic oxygen and ultraviolet radiation. Patent applications have been submitted for these concepts.

The Space Constructible Radiator (SCR) Life Test heat pipe performance testing is currently conducted at NASA/Johnson Space Center as part of the Advanced Technology Development Program. The SCR is a dual passage, monogroove heat pipe radiator designed and manufactured by Grumman Aerospace for NASA. The heat pipe has many aerospace applications since it can transport a large amount of heat with a compact lightweight design. As the micro-meteoroid/orbital debris environment worsens, it may be advantageous to add the heat pipe radiator to the Space Station's thermal control system. The SCR Life Test has been operating over the last 10 years and will continue until the year 2000. The overall heat transfer coefficient has decreased from 792 W/K (1500 Btu/Hr-°F) to 475 W/K (900 Btu/Hr-°F) but appears to have stabilized. This paper summarizes the SCR Life Test setup and the test results to date.

This paper defines the value of free-flying cameras to the Space Station. The use of free-flying cameras is an alternative to reliance on fixed cameras. The analysis is based upon results from recent neutral buoyancy evaluations of a free-flying camera known as the Supplemental Camera and Maneuvering Platform (SCAMP). SCAMP was evaluated for inspection and viewing capabilities that will be required by Space Station. Test results demonstrated that a free-flying camera could be used effectively for inspecting structure, viewing labels, providing views for control of extravehicular robotics (EVR) and for ground assistance during extravehicular activity (EVA) tasks.

The size and complexity of Space Station has driven the need for an accurate, reliable analytical tool to assess the extravehicular activity (EVA) crew interfaces at the worksite. On previous spacecraft, each worksite was developed and validated through Neutral Buoyancy underwater testing by the crew using mockups. For spacecraft requiring a significant amount of EVA over large areas, like Space Station, the cost of conducting underwater tests for each of the many hundred worksites becomes prohibitive. Therefore, limited testing must be augmented by accurate graphical analysis. The Unigraphics II, which is the Computer Aided Design (CAD) system for the International Space Station Alpha (ISSA) Product Group 1 design, was selected and developed. It has a major advantage of easy and rapid access to the accurate and updated Space Station design. The design can be rapidly obtained electronically from layouts, detail drawings, assembly drawings or the Electronic Development Fixture (EDF).

The objective of this paper is to present results of MDA's payload vibration isolation system research and development program. A unique isolation system with passive or active capabilities designed to provide isolation down to 10-6 g was developed and tested in our 1-g testbed under simulated microgravity conditions. Fluid and electrical umbilicals are also included in the system. The established isolation system performance requirements were met and the testbed data were used to refine our analytical models for predicting flight performance. Simulations using an updated Space Station configuration showed that the payload microgravity requirement can be met by upgrading the hardware from laboratory to flight tolerances and improving the control system design. The next step is to flight test the systems verified in 1 g on the STS/SPACEHAB using a middeck locker size development unit.

Future military aircraft will demand lower cost and lower weight subsystems that are more reliable, and easier to maintain and support. To identify and develop subsystems integration technologies that could provide benefits such as these to current and future military aircraft, the Air Force Wright Laboratory (WL/FIVE) initiated the Subsystem Integration Technology (SUIT) program in 1991. McDonnell Douglas Aerospace (MDA) together with Pratt and Whitney (PWA), and AlliedSignal Aerospace Systems and Equipment (ASE) was one of three teams that participated in Phase I of the SUIT program. The MDA Team's goal was to conceptually formulate a SUIT approach which would provide significantly reduced weight and costs while increasing cooling and power generation capabilities. These goals were achieved with a new and innovative energy subsystem suite which integrates aircraft and engine subsystem power, cooling, pumping, and controls.

The unique development environment of NASA's Space Station Freedom (SSF) creates numerous challenges to the design of a common user interface for operating the spacecraft. Astronauts on board SSF will utilize multi-purpose workstations as their primary command and control interface to the vehicle. With the exception of some dedicated hardware controls, the vast majority of the workstation user interface will be implemented in software. The specification and design of the SSF user interface requires the synthesis of on-orbit operational requirements with multiple systems' functional requirements, all of which emanate from geographically and organizationally distributed entities. Human factors requirements as well as constraints imposed by the SSF Displays and Controls (D&C) system architecture are additional considerations.

As manned spacecraft have evolved into larger and more complex configurations, the mandate for preventing, detecting, and extinguishing on-board fires has grown proportionately to ensure the success of progressively ambitious missions. The closed environment and high value of manned spacecraft offer the Fire Detection and Suppression (FDS) systems designer significant challenges. With the presence of Oxygen (O2), flammable materials, and ignition sources, it is impossible to completely remove the likelihood of a spacecraft fire. Manned spacecraft contain these three ingredients for fire; therefore, it becomes profitable to review past designs of FDS systems and ground testing to determine system performance and lessons learned in the past for present and future applications.

The Space Station Freedom program requires long-life thermal control coatings that are stable in low Earth orbit (LEO). To provide designers with a variety of coatings and optical properties, improvements were made to existing coatings, and new thermal control coatings were developed. Anodized aluminum was demonstrated to be an acceptable substrate for inorganic thermal control coatings such as Z-93. Mixtures of Z-93 with stable black oxides provided a wide range of optical properties and were stable in a simulated LEO environment. In addition, sulfuric acid anodized aluminum was developed to a production status to provide controlled optical properties for many aluminum alloys.

The National Aeronautics and Space Administration (NASA) Space Station Freedom (SSF) will use Flow-Through Radiators (FTRs) to reject waste heat that is collected from the on-board Heat Acquisition Devices (HADs). The waste heat is sent to the FTRs via the Pump Module Assembly (PMA) subsystem of the External Active Thermal Control System (EATCS). Two developmental FTR panels were integrated with the EATCS Ground Test Article (GTA). The integrated components were investigated under a thermal/vacuum environment in Thermal/Vacuum Chamber A at NASA/JSC during June, 1992. A detailed SINDA/FLUINT FTR model was developed to predict the steady-state thermal/hydraulic performance of the FTRs. A simplified SINDA/FLUINT FTR model was also developed for use in the GTA integrated model. Schematics and plots comparing the test data and model results are presented for both steady-state and transient conditions.

This paper presents an overview of the design and operation of the internal thermal control system (ITCS) developed for Space Station Freedom by the NASA-Johnson Space Center and McDonnell Douglas Aerospace to provide cooling for the resource nodes, airlock, and pressurized logistics modules. The ITCS collects, transports, and rejects waste heat from these modules by a dual-loop, single-phase water cooling system. ITCS performance, cooling, and flow rate requirements are presented. An ITCS fluid schematic is shown and an overview of the current baseline system design and its operation is presented. Assembly sequence of the ITCS is explained as its configuration develops from Man Tended Capability (MTC), for which node 2 alone is cooled, to Permanently Manned Capability (PMC) where the airlock, a pressurized logistics module, and node 1 are cooled, in addition to node 2.