SAWE Technical Papers

@inproceedings{3670,

title = {3670. Soil Moisture Active Passive Mission Mass Properties Optimization and Analysis},

author = {Morgan Hendry IV and Erich Slimko},

url = {https://www.sawe.org/product/paper-3670},

year  = {2016},

date = {2016-05-01},

booktitle = {75th Annual Conference, Denver, Colorado},

pages = {22},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Denver, Colorado},

abstract = {The Soil Moisture Active Passive (SMAP) mission creates global soil moisture measurements every 2-3 days by spinning a lightweight, ~6 meter diameter deployable mesh reflector. This represents the largest spinning antenna ever to be used on a NASA satellite. The paper discusses the mass properties challenges that were overcome in implementing the mission, including the definition of a combined imbalance term called the Effective Product of Inertia, the selection of an optimum Spun Instrument focal length and edge offset that fostered a reduced mass properties sensitivity, the inclusion of late game Spun Instrument configuration flexibility, the use of balance mass locations that leveraged the large moment arm provided by the reflector, and a subassembly measurement campaign that reduced mass properties uncertainties on testable components.},

keywords = {18. Weight Engineering - Spacecraft Design, 34. Advanced Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3619,

title = {3619. Inflatable Antenna For CubeSat: Design, Fabrication, Deployment And Tests},

author = {A. Babuscia},

url = {https://www.sawe.org/product/paper-3619},

year  = {2014},

date = {2014-05-01},

booktitle = {73rd Annual Conference, Long Beach, California},

pages = {13},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Long Beach, California},

abstract = {CubeSats and small satellites have potential to provide means to explore space and to perform science in a more affordable way. As the goals for these spacecraft become more ambitious in space exploration, moving from Low Earth Orbit (LEO) to Geostationary Earth Orbit (GEO) or further, the communication systems currently implemented will not be able to support those missions. One of the bottlenecks is the antennas' size, due to the close relation between antenna gain and dimensions. Current antennas for CubeSats are mostly dipole or patch antennas with limited gain. Deployable (not inflatable) antennas for CubeSats are currently being investigated, but these solutions are affected by the challenge of packaging the whole deployable structure in a small spacecraft. 

The work that we propose represents the first attempt to develop an inflatable antenna for CubeSats. Inflatable structures and antennas can be packaged efficiently, occupying a small amount of space, and they can provide, once deployed, large dish dimension and correspondent gain. Inflatable antennas have been previously tested in space (Inflatable Antenna Experiment, STS-77). However they have never been developed for small spacecraft such as CubeSats, where the packaging efficiency, the deployment, and the inflation represent a challenge. 

The article covers the design and radiation model for the antenna. Details of the antenna's fabrication and related issues are illustrated as well as the mechanism to fold and deploy the antenna in space. The results of the experimental tests are described. The extension of simulation to cover the X-Band is also discussed. Future work in the development of the antenna will include the improvement of the fabrication process and the design of a 3U CubeSat mission to be proposed as a technical demonstration.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3593,

title = {3593. Trade Study of System Level Ranked Radiation Protection Concepts for Deep Space Exploration},

author = {Jeffrey Cerro},

url = {https://www.sawe.org/product/paper-3593},

year  = {2013},

date = {2013-05-01},

booktitle = {72nd Annual Conference, St. Louis, Missouri},

pages = {18},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Saint Louis, Missouri},

abstract = {A strategic focus area for NASA is to pursue the development of technologies which support exploration in space beyond the current inhabited region of low earth orbit. An unresolved issue for crewed deep space exploration involves limiting crew radiation exposure to below acceptable levels, considering both solar particle events and galactic cosmic ray contributions to dosage. Galactic cosmic ray mitigation is not addressed in this paper, but by addressing credible, easily implemented, and mass efficient solutions for the possibility of solar particle events, additional margin is provided that can be used for cosmic ray dose accumulation. As a result, NASA's Advanced Engineering Systems project office initiated this Radiation Storm Shelter design activity. This paper reports on the first year results of an expected 3 year Storm Shelter study effort which will mature concepts and operational scenarios that protect exploration astronauts from solar particle radiation events. Large trade space definition, candidate concept ranking, and a planned demonstration comprised the majority of FY12 activities. A system key performance parameter is minimization of the required increase in mass needed to provide a safe environment. Total system mass along with operational assessments and other defined protection system metrics provide the guiding metrics to proceed with concept developments. After a downselect to four primary methods, the concepts were analyzed for dosage severity and the amount of shielding mass necessary to bring dosage to acceptable values. Besides analytical assessments, subscale models of several concepts and one full scale concept demonstrator were created. FY12 work terminated with a plan to demonstrate test articles of two selected approaches. The process of arriving at these selections and their current envisioned implementation are presented in this paper.},

keywords = {18. Weight Engineering - Spacecraft Design, 19. Weight Engineering - Spacecraft Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3567,

title = {3567. Crew Transfer Options for Servicing of Geostationary Satellites},

author = {Jeffrey Cerro},

url = {https://www.sawe.org/product/paper-3567},

year  = {2012},

date = {2012-05-01},

booktitle = {71st Annual Conference, Bad Gögging, Germany},

pages = {28},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Bad Gögging, Germany},

abstract = {In 2011, NASA and DARPA undertook a study to examine capabilities and system architecture options which could be used to provide manned servicing of satellites in Geostationary Earth Orbit (GEO). The study focused on understanding the generic nature of the problem and examining technology requirements, it was not for the purpose of proposing or justifying particular solutions. A portion of this study focused on assessing possible capabilities to efficiently transfer crew between Earth, Low Earth Orbit (LEO), and GEO satellite servicing locations. This report summarizes the crew transfer aspects of manned GEO satellite servicing. Direct placement of crew via capsule vehicles was compared to concepts of operation which divided crew transfer into multiple legs, first between earth and LEO and second between LEO and GEO. In space maneuvering via purely propulsive means was compared to in-space maneuvering which utilized aerobraking maneuvers for return to LEO from GEO. LEO waypoint locations such as equatorial, Kennedy Space Center, and International Space Station inclinations were compared. A discussion of operational concepts is followed by a discussion of appropriate areas for technology development.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3467,

title = {3467. Weight Analytics},

author = {Chuck Higgins},

url = {https://www.sawe.org/product/paper-3467},

year  = {2009},

date = {2009-05-01},

booktitle = {68th Annual Conference, Wichita, Kansas},

pages = {13},

address = {Wichita, Kansas},

abstract = {Weight Analytics addresses the gathering, analyzing, reporting and sharing of critical weight data and opening new horizons for the use of this knowledge within the design process. In this paper we will show the use of a cubical data base, allowing the user to look at an infinite number of dimensions of data quickly, easily and intuitively. Automation of the collection, storage, analysis and reporting on this disparate data has until now been challenging at the least. With the flexibility of the Rubix Cube, data is quickly and easily sliced and diced along any number of 

relationships thus 'elevating data to knowledge.'},

keywords = {10. Weight Engineering - Aircraft Design, 18. Weight Engineering - Spacecraft Design, 22. Weight Engineering - Structural Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3481,

title = {3481. Flight Test Analysis of LOX/Propylene Upper Stage Engine},

author = {Kay Gemba and Deepak Verma},

url = {https://www.sawe.org/product/paper-3481},

year  = {2009},

date = {2009-05-01},

booktitle = {68th Annual Conference, Wichita, Kansas},

pages = {17},

address = {Wichita, Kansas},

abstract = {The objective of this paper is to present testing and analysis of an early prototype upper stage engine which could be optimized and evolved into a second stage engine for a Nanosat Launch Vehicle (NLV). The NLV is designed to deliver a nominal 10 kg payload to LEO and is being developed by the California Launch Vehicle Education Initiative (CALVEIN), a partnership program between Garvey Spacecraft Corporation (GSC) and California State University, Long Beach (CSULB) [3]. 

The engine is pressure-fed and uses LOX/propylene as propellants. It is designed to operate at a chamber pressure of 1 MPa and provide a vacuum thrust of 2000 N. Propylene was chosen as a propellant because it provides a higher specific impulse than RP-1 with comparable density at cryogenic temperatures [16]. 

This paper presents a first iteration of the preliminary design intended for space operations with an expansion ratio of 70 as well as the testing of its sea level version with an expansion ratio of 4. The space engine is designed with targeted combustion efficiency of 95% and nozzle efficiency of 98%, corresponding to a specific impulse of 347 s. 

A static fire test of the engine, shown in Figure 9, with expansion ratio of 4 has been conducted twice at sea level with a burn time of 15 seconds and 5 seconds, respectively. A flight test, shown in Figure 11, has also been conducted to test the capabilities of the engine. Recorded data will be used to assess previous assumed efficiencies and refine, if necessary, the shape of the nozzle and configuration of the injector. The next steps include implementing necessary changes to the engine to achieve better performance for future testing. This paper will also address ways to manage and reduce overall engine weight to improve performance.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3457,

title = {3457. How Mass Properties Affect Satellite Attitude Control},

author = {Richard Boynton},

url = {https://www.sawe.org/product/paper-3457},

year  = {2008},

date = {2008-05-01},

booktitle = {67th Annual Conference, Seattle, Washington},

pages = {21},

address = {Seattle, Washington},

abstract = {The success of a satellite mission is highly dependent on the accuracy of the measurement of its mass properties before flight and the proper ballasting of the satellite to bring the mass properties within tight limits. Failure to properly control mass properties can result in the satellite tumbling end over end after launch, or quickly using up its thruster capacity in an attempt to point in the correct direction. Solar panels must continue to point toward the sun as the satellite orbits the earth. Telescopes must point earthward. Satellite attitude control systems generally consist of a closed loop of measurement and correction of the spacecraft's attitude such that it is constantly driven into its desired nominal orientation, effectively rejecting any disturbances imposed on the satellite, such as variations in the earth's magnetic field, nonspherical shape of the Earth, lunar and solar perturbations, drag of the residual atmosphere on the solar array, and solar radiation pressure, or by movement of mechanical parts within the satellite. This paper discusses the different means of attitude control: thrusters, momentum wheel, spin stabilization, gravity gradient stabilization, and magnetic field control, with emphasis on the relationship of mass properties to these control methods.},

keywords = {06. Inertia Measurements, 18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3460,

title = {3460. Using a Two-Plane Spin Balance Instrument to Balance a Satellite Rotor About Its Own Bearings},

author = {Paul Kennedy and Daniel Otlowski and Brandon Rathbun and Kurt Wiener},

url = {https://www.sawe.org/product/paper-3460},

year  = {2008},

date = {2008-05-01},

booktitle = {67th Annual Conference, Seattle, Washington},

pages = {21},

address = {Seattle, Washington},

abstract = {This paper addresses the problem of statically and dynamically balancing a satellite, mounted antenna rotor supported on its own bearings, and driven by a motor in the satellite body. The satellite body is considered a stationary platform, (stator) for this procedure and is not part of the balancing problem. The antenna rotor is isolated and balanced independently while spinning on its own bearings. In order to measure the unbalance, a method is developed to utilize a two-plane vertical axis spin balance machine. Rather than using the gas bearing rotor of the measuring instrument and spinning the entire satellite, the satellite body (stator) is attached to the balancing machine table, which is held stationary, and the satellite 'rotor' is spun on its own bearings. Forces due to the unbalance are measured by the Spin Balance Machine force transducers. The method is compared to a similar procedure using a single plane spin balancer and to methods using 'work reversal' methods to balance the rotor by spinning the entire satellite. The accuracy of this procedure is compared to the basic balance capability of the spin balance instrument when used in the conventional manner.},

note = {Mike Hackney Best Paper Award},

keywords = {06. Inertia Measurements, 18. Weight Engineering - Spacecraft Design, Mike Hackney Best Paper Award},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3386,

title = {3386. Design of a Lunar Lander for Concept Exploration and Refinement},

author = {Leo Bocam},

url = {https://www.sawe.org/product/paper-3386},

year  = {2006},

date = {2006-05-01},

booktitle = {65th Annual Conference, Valencia, California},

pages = {14},

publisher = {Society of Allied Weight Engineers},

address = {Valencia, California},

organization = {Society of Allied Weight Engineers},

abstract = {Orbital Sciences Corporation developed reference missions, system requirements, and exploration architecture concepts in support of the United States Vision for Space Exploration. These efforts were performed as part of the National Aeronautics and Space Administration (NASA) Concept Exploration and Refinement (CE&R) study, which examined initial missions to Low Earth Orbit (LEO), initial and sustaining missions to the Moon, and initial and sustaining missions to Mars. As in all aerospace development efforts, safety, cost, and schedule were the driving factors. Orbital focused first on developing a sustainable and affordable LEO and lunar transportation architecture. The first steps for executing a lunar mission include placing the astronauts into LEO, transporting them to Low Lunar Orbit (LLO), and eventually to the lunar surface with extended mission duration capability. The key element of the lunar transportation architecture is the lunar lander, which provides transportation to and from the lunar surface. Lunar lander configurations assessed included single stage, two stage, reusable, and expendable, which were all examined with different propellant combinations. Orbital received Level 1 requirements from NASA and decomposed them into requirements specific to the lunar lander based on architectural level trade studies performed during the contract. The iterative process between design and systems engineering ultimately allowed Orbital to recommend a lunar lander design to NASA. Orbital?s solution was optimized into a single stage reusable Human Lunar Lander (HLL) with full abort coverage to and from the lunar surface. In addition, a cargo variant known as the Cargo Lunar Lander (CLL) was developed with an emphasis on maximizing commonality and reducing cost. The resulting design was optimized using mass properties estimation and performance analysis. The recommended HLL possesses excellent performance capability and safety characteristics, such as abort capability. This paper discusses the assumptions and design considerations Orbital used to develop the HLL design concept, as well as the critical crew safety features of the configuration.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3323,

title = {3323. Optimizing Composite Rocket Motor Development Using Advanced Evolutionary Algorithms},

author = {Frank Abdi and Dr. Myles L. Baker},

url = {https://www.sawe.org/product/paper-3323},

year  = {2004},

date = {2004-05-01},

booktitle = {63rd Annual Conference, Newport, California},

pages = {13},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Newport, California},

abstract = {The design of composite motor cases for solid rockets is a very complex undertaking, and requires the careful consideration of many factors. One of the most obvious challenges is ensuring that the materials chosen, the winding sequence, and the orientation and ordering of the winding plies results in a pressure vessel with the highest possible strength for a given weight. This is significantly complicated by the fact that composite motor cases are not simple pressurized bottles, but have stringent requirements for integration with the rest of the vehicle in the form of skirts, nozzles, fins, lugs, etc. At the skirts, sufficient compliance must be designed into the structure so as not to cause local failures when the pressure vessel expands under pressure, but sufficient strength must be available to carry the required axial, bending, and transverse loads. At the pole pieces, the bond between the composite case material and the typically metallic pole hardware must be able to withstand the stresses from differential thermal expansion as well as mechanical stresses due to pressurization. The manufacturing and life prediction process has been applied to the analysis and optimization of a composite motor case with excellent results.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3324,

title = {3324. One Idea for a Next Generation Space Shuttle},

author = {Ian O. MacConochie and Jeffrey Cerro},

url = {https://www.sawe.org/product/paper-3324},

year  = {2004},

date = {2004-05-01},

booktitle = {63rd Annual Conference, Newport, California},

pages = {14},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Newport, California},

abstract = {In this configuration, the current Shuttle external tank serves as core structure for a fully reusable second stage. This stage is equipped with wings, vertical fin, landing gear, and thermal protection. The stage is geometrically identical to (but smaller than) a single stage that has been tested hypersonically, supersonically, and subsonically in the NASA Langley Research Center wind tunnels. The three LOX/LH engines that currently serve as main propulsion for the Shuttle Orbiter, serve as main propulsion on the new stage. The new stage is unmanned but is equipped with the avionics needed for automatic maneuvering on orbit and for landing on a runway. Three rails are installed along the top surface of the vehicle for attachment of various payloads. Payloads might include third stages with satellites attached, personnel pods, propellants, or other items.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3326,

title = {3326. Evolution of a Blended Lifting Body for the Orbital Space Plane},

author = {Ryan D. Leo},

url = {https://www.sawe.org/product/paper-3326},

year  = {2004},

date = {2004-05-01},

booktitle = {63rd Annual Conference, Newport, California},

pages = {12},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Newport, California},

abstract = {NASA?s Orbital Space Plane (OSP) program incorporated elements of past efforts such as the Space Transportation Architecture Study (STAS), Crew Return Vehicle (CRV), and Space Launch Initiative (SLI). Significant goals of these efforts were to improve safety, reduce the cost of crew transportation to space, and to provide robust crew rescue and crew transfer capability for the International Space Station (ISS). NASA?s requirements for OSP were assessed and evaluated, and several critical design drivers were derived: number of crew, mission duration, launch vehicle throw weight, vehicle lift efficiency, flight rate, reusability, and ascent abort and emergency return capabilities. NASA Langley Research Center?s HL-20 shape was used as an initial reference design to better understand and assess the impact of NASA requirements as they were flowed down to the vehicle subsystem level. Configuration and aerodynamic trades were conducted to optimize the performance of the OSP in response to these requirements. Most importantly, issues of volumetric efficiency, high L/D for cross range, low wing loading for reduced landing speed, and passive stability for all abort conditions were addressed. As the optimization process continued, the HL-20 initial reference shape eventually evolved into the Blended Lifting Body (BLB). The BLB combines volumetric efficiency with superior aerodynamic qualities and was designed to launch vertically and land horizontally. The BLB design offers an optimized configuration with excellent aerodynamic performance and may have many other flight applications. This paper discusses the evolution process, design solutions and features of the configuration used during the development of the BLB for the OSP program.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3332,

title = {3332. Generic Space Shuttle Program (SSP) Flight Feasibility Assessment Techniques},

author = {Ross A. Abotteen},

url = {https://www.sawe.org/product/paper-3332},

year  = {2004},

date = {2004-05-01},

booktitle = {63rd Annual Conference, Newport, California},

pages = {13},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Newport, California},

abstract = {The SSP at National Aeronautics and Space Administration (NASA)/Johnson Space Center (JSC) employs these Generic Flight Feasibility Assessment Techniques (FFAT) presented in this paper. The FFAT results are used for flight baseline, ?what-if?/?trade-off? flight analysis, and flight planning of the Shuttle manifest. The FFAT are early analysis tools in the SSP. These early analysis results are refined using NASA Flight Ops simulators. The FFAT process starts by selecting a flight from the SSP Flight Assignment Working Group (FAWG) Planning Manifest referred to as Flight Under Design (FUD). This flight would have an Orbiting Vehicle assignment (OV-103 Discovery, or OV-104 Atlantis, or OV-105 Endeavour). It would also have a Payload Bay (PLB) manifested complement of payloads and a projected Launch Date. A historically flown Shuttle mission is selected from the SSP database. It is referred to as Reference Flight (RF). The RF mission parameters are very similar to the FUD parameters. The RF Main Engine Cut-Off (MECO) Mass Lift Capability is compared to the FUD flight MECO Mass requirements in order to determine if the flight can lift such Mass requirements. In effect, the Ascent Performance Margin (APM) in this comparison must be positive and meeting/exceeding SSP Manager?s Reserve Policy value. All applicable flight Ascent Performance Partials (APP) are considered in this comparison for determining the APM.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3304,

title = {3304. The Many Options For Future Earth-to-Orbit Transports},

author = {Ian O. MacConochie},

url = {https://www.sawe.org/product/paper-3304},

year  = {2003},

date = {2003-05-01},

booktitle = {62nd Annual Conference, New Haven, Connecticut},

pages = {10},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {New Haven, Connecticut},

abstract = {There are many options available for earth-to-orbit transportation systems. The shuttle has made over 100 trips to orbit serving on occasion as a mini-space station but principally for cargo delivery and crew transfer. Its cargo volume and shape is so large that it was possible to deliver to orbit a large telescope (the Hubble). In addition to the shuttle, there are many excellent expendable multi-stage vehicles available. The next progression in space transportation might involve the use of an array of vehicles with greater reusability of elements. Invoking reusability, however, usually leads to greater development costs but hopefully lower operating costs. A number of approaches to launch vehicle design are shown including several with expendable upper stages but a reusable first stage; also shown are several concepts involving the use of both reusable first and second stages.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3305,

title = {3305. The Good and Bad about Single-Stage-to Orbit},

author = {Ian O. MacConochie},

url = {https://www.sawe.org/product/paper-3305},

year  = {2003},

date = {2003-05-01},

booktitle = {62nd Annual Conference, New Haven, Connecticut},

pages = {12},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {New Haven, Connecticut},

abstract = {The good aspect of single-stage-to-orbit is that there are no lower stages that must be recovered. If the launch system is multi-stage, recovery is made difficult because staging at any reasonable lapsed time from liftoff means that the lower stages are a long way from the launch site. When the reusability rule is invoked, early stages must be equipped with wings and wheels, parachutes, or some other means of returning the stage to the launch site. Disadvantages of single stage include high sensitivity to weight growth, difficulty in obtaining meaningful payloads for reasonable assumptions of technology, and tendency for center of gravity to migrate rearward as structural weight is reduced and main propulsion system weight is increased. These issues are discussed in the paper.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3286,

title = {3286. Developing Metal Matrix Composites For Ordnance and Aerospace Applications},

author = {Brian L. Gordon},

url = {https://www.sawe.org/product/paper-3286},

year  = {2002},

date = {2002-05-01},

booktitle = {61st Annual Conference, Virginia Beach, Virginia, May 18-22},

pages = {10},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Virginia Beach, Virginia},

abstract = {Continuous-fiber reinforced metal matrix composites (MMCs) present significant improvements in specific strength and specific stiffness over conventional monolithic alloys. MMC materials offer tremendous potential for U.S. Army applications such as lightweight projectiles and gun systems. For advanced artillery projectiles, MMCs can be used to manufacture shell bodies 50% lighter than steel shells with more payload carrying capacity. MMCs could also play an important part in the development of lightweight jackets for steel-lined gun barrels, which are longer and stiffer, with higher-pressure capability at equivalent weights compared to current designs. In addition, NASA has repeatedly identified metal matrix composite development and lightweight cryogenic compatible materials as high priorities for supporting advanced space transportation requirements. MMC technology could have direct application to cryogenic-compatible propulsion system feed lines and ducts, and could be expanded with further investment to be applied to cryogenic tanks and primary structure as well. Unfortunately, manufacturing difficulties have limited the number of applications and prevented these materials from fully penetrating the marketplace. To address these technology needs and manufacturing shortfalls, Touchstone is developing a new material system and manufacturing method in which the advantages of MMCs are married with the manufacturing techniques of PMCs. This technology demonstrates that MMC structures can be produced in an analogous manner to PMCs utilizing traditional tape/fiber placement processing techniques. Additionally, the process represents both on-the-fly and out-of-autoclave composites technology with potentially large cost savings. This paper describes Touchstone?s current efforts to advance this unique technology.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3117,

title = {3117. The Development of Sea Launch Mass Properties},

author = {Cannon and Wetzel},

url = {https://www.sawe.org/product/paper-3117},

year  = {2001},

date = {2001-05-01},

booktitle = {60th Annual Conference, Arlington, Texas, May 19-23},

pages = {15},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Arlington, Texas},

abstract = {The Sea Launch program is a unique international collaboration to provide commercial launch services using hardware made in the United States, Russia, and the Ukraine. The chronology of agreements that led to a successful method for calculating mass properties is presented, reconciling diverse, culturally rooted philosophies of weights engineering. Additionally, novel methods for verification of weight and center of mass are discussed, based on a minimum cost, short-cycle commercial approach to development and operations. Lessons learned for future programs include recognizing and using best practices from multiple sources to create a hybrid, program-unique approach to mass properties management.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{2457,

title = {2457. Balancing Lunar Prospector},

author = {A Tilley},

url = {https://www.sawe.org/product/paper-2457},

year  = {1999},

date = {1999-05-01},

booktitle = {58th Annual Conference, San Jose, California, May 24-26},

pages = {39},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {San Jose, California},

abstract = {This paper is a narrative of Lunar Prospector design evolution, the inherent mass property problems encountered and the extensive measurement and ballasting program which was needed to verify the flight readiness of the vehicle. Lunar Prospector was a spin stabilized lunar research satellite launched from a spinning upper stage. Economic considerations ruled out sophisticated control systems and compelled the system design to be simple and naturally balanced; dynamic balance requirements for spin stabilization drove the need for a high degree of precision in the final results. Nine of the ten vehicle mass properties (mass, three centers of mass, three moments of inertia and two of the three products of inertia) were controlled parameters.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{2481,

title = {2481. X-34 Reusable Launch Vehicle Technology Demonstrator Mass Properties Control},

author = {K Bocam},

url = {https://www.sawe.org/product/paper-2481},

year  = {1999},

date = {1999-05-01},

booktitle = {58th Annual Conference, San Jose, California, May 24-26},

pages = {18},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {San Jose, California},

abstract = {(None - PRESENTATION)},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{2411,

title = {2411. X-34 Reusable Launch Vehicle Technology Demonstrator Mass Properties Engineering Challenges},

author = {K J Bocam},

url = {https://www.sawe.org/product/paper-2411},

year  = {1998},

date = {1998-05-01},

booktitle = {57th Annual Conference, Wichita, Kansas, May 18-20},

pages = {13},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Wichita, Kansas},

abstract = {X-34 is an Air Launched Reusable Hypersonic Testbed and a key component in NASA's Reusable Launch Vehicle (RLV) Program. This paper examines the mass properties related challenges resulting from developing a reusable hypersonic technology demonstrator in a concurrent engineering environment. An overview of the scope and content of the X-34 Program and its role in NASA's RLV Program is given. Unique elements of developing a low cost RLV demonstrator and their implications to vehicle mass properties are addressed.},

keywords = {18. Weight Engineering - Spacecraft Design},

pubstate = {published},

tppubtype = {inproceedings}

}