Topics/Technology Area: Spacecraft / Solar Sails
Program Manager: Edward (Sandy) Montgomery - NASA/MSFC

Three Selections:

1. Able Engineering Company - Scalable Solar Sail System Development and Ground Demonstration

The ABLE-SRS team proposes a disciplined systems engineering development program that culminates in a ¼-scale demonstration of a 100-m class solar sail system. ABLE's kinematically determinant large-system design and test approach produces a solar sail system amenable to credible flight qualification via practical ground testing. This assertion is backed by ABLE's 100% success rate in developing and ground-test-qualifying very large space structures for flight. Our proposed design and system demonstration test approach will prove solar sails are ready for Sentinel class (100-m) and future larger missions. We propose a phased technology validation effort begins in phase one with detailed design studies to identify technical issues arising at the system level. In phase two we develop detailed designs, validate analytical models and begin validating selected element and subassembly hardware. In phase three we integrate our validated subassemblies into a sail system, and demonstrate the flight-worthiness of our sail system design approach with detailed and system-wide ground tests. Under the systems engineering leadership of ABLE Engineering, our team joins leading NASA and industry solar sailing technologists, flight-seasoned engineers, and private consultants to focus a wealth of aerospace engineering and solar sailing program experience on the solar sail system demonstration task. The program, over three phases, builds a solid engineering basis for the solar sail system demonstrator. A detailed series of tests are performed to demonstrate and measure performance, and correlations with pre-test modeling validate our tools for future missions. Our ground-testable sail system will meet required sail performance with good margins and high reliability. Our system design and demonstration of this technology under NASA funding will ready sail technology not only for the critical Sentinel missions, but also for a series of progressively more challenging missions planned by NASA, NOAA, and the DoD.

2. L’Garde Inc. - Development of an ultra-light weight and eminently scalable inflatable deployed and supported Solar Sail Subsystem

L'Garde, Inc. shall use its space-flight experience in inflatable, rigidizable and membrane space structures to develop a scalable, lightweight and cost-effective solar sail subsystem to TRL-6. We propose a carefully planned program, based on Solar Sail Design, Manufacturing and Tests we have done since 1994, and are doing now for the Team Encounter mission. We have also brought to the effort a strong team of experts in (a) Structures, Materials and Dynamic testing (NASA-LaRC), (b) Mission/Orbit Definition, Systems Engineering and GNC (JPL and Ball Aerospace); and (c) Team Encounter for the Outreach Program. The NASA requirements and the systems analysis results will flow down to the design analyses of the Solar Sentinel Solar Sail subsystem concept design. This will result, after the scaling analyses, to the design of three ground-test items proposed to bring the Solar Sail SOA to TRL-6: The first will be a sub-scale, 10mX10m test unit for ambient deployment and design shake-down tests. The second will also be a 10mX10m unit for vacuum deployment, rigidization and dynamics testing. Based on the results of these two tests, the third test item, the 20mX20m full-scale test item design will be updated. The proposed effort culminates in the fabrication and test of the 20X20m unit at the NASA/GRC's Plum Brook facility. This final test unit shall be shown to meet or exceed all solar sail requirements by analysis and test results, followed by the Solar Sentinel Mission Concept design update. The proposed solar sail structure consists of four ultra-light, inflatable-rigidizable booms in a cruciform configuration, controllably deployed within a narrow geometric envelope, via a patented L'Garde process. The proposed rigidization is the Sub-Tg (or cold) rigidization method. The proposed sail material is 2-micron Mylar, metalized on both sides, proven by space-environment analyses and test to amply withstand the environments of most solar sail missions. L'Garde, however, is totally open to the use of any NASA-preferred materials. The sail will be reinforced against tearing and tear propagation and will be attached to the cruciform structure in a following-load configuration. This results in the lightest possible scalable structure and virtually eliminates bending loads and long-column buckling. Based on our analytical and experimental results to date, we believe this is the most efficient path to an operational solar sail for NASA.

3. JPL - Solar Sails GNC Took Kit

The technology needed to accurately predict realistic flight performance estimates for solar sail missions does not yet exist. A team comprised of technical experts from the Jet Propulsion Laboratory, California Institute of Technology and Ball Aerospace & Technologies Corp. in Boulder, Colorado proposes to develop an integrated set of simulation tools to predict, re-calibrate, and optimize the trajectory, maneuvers, and propulsive performance of a sail during a representative flight mission. These trajectory optimization and sail-craft guidance, navigation and control (GNC) simulation tools are necessary for realistic studies of solar sail missions. Additional professional participants from government, industry and universities will add strategic expertise in modeling and simulation. This team is uniquely capable of developing a realistic solar sail GNC toolset owing to their actual low-thrust flight experience on New Millennium DS1 and development of a commercial sail-craft. The toolset will advance the technology readiness level to that needed for rapid design and planning of solar sail flight mission concepts. In addition, the toolset will be designed so that it can be easily integrated into an optimal GNC system for the flight of future sail missions. The toolset will include analytical models for (1) solar radiation pressure acting on the sail, (2) disturbance forces acting on the sail from gravitation and thermal torques, (3) orbital mechanics, (4) sail structural dynamics, (5) attitude control system dynamics, (6) navigation sensors, and (7) environmentally induced changes to sail properties. A mission concept from the class of Solar Wind Sentinels will also be modeled, and the integrated simulation tools will be demonstrated on that candidate mission.

Topics/Technology Area: Spacecraft / Aerocapture
Program Manager: Bonnie James - NASA/MSFC

Six Selections:

1. Langley Research Center - High Temperature Composites and Adhesives for Reduced-Mass Aeroshells

The objective of the work proposed here is to leverage several emerging structures and materials technologies to significantly reduce the mass of aeroshell systems. Because reduced mass translates directly into larger payloads and increased science, the importance of accomplishing this goal is common to nearly all planetary missions that utilize aerocapture. Current aeroshell designs typically use metallic structure. Sometimes conventional composite materials are incorporated into the design. The amount of thermal protection required for these materials is relatively great, and the resulting aeroshells are relatively massive. There are now new families of high-temperature resins and composite material manufacturing techniques that have the potential to significantly reduce mass and improve aeroshell design. The proposed effort is intended to demonstrate that the use of these technologies can result in significant mass reductions (approaching 30%) for the primary aeroshell structure and its associated thermal protection system (TPS). An additional goal is to show that these mass reductions can be obtained at relatively low cost, and without the high risk often associated with alternative aeroshell designs or planetary entry systems. The program will determine optimized aeroshell system masses for a range of planetary entry environments (including Titan and Neptune) with acceleration loads of 3-20g and peak structural temperatures ranging from 400-700 degrees Fahrenheit (for the primary structure that lies behind the TPS). The optimized systems should result in a significant mass savings for relatively less challenging missions (such as Titan aerocapture), and will hopefully go further to become enabling for more difficult missions (such as Neptune aerocapture). The approach will be to determine optimized aeroshell concepts through analysis of several exploration mission scenarios. State-of-the art high-temperature composite materials, resins, and adhesives will then be examined for incorporation into these concepts. The program will verify fabrication methods for lightweight compound curvature aeroshells through the fabrication of representative coupon samples, structural elements, and finally system level subcomponents, all of which will be tested in relevant environments. This effort will be a natural extension of work that has already been done to incorporate high-temperature materials into NASA aircraft and reusable launch vehicle (RLV) programs. The effort will build on the heritage of existing aeroshell design, and will use a building-block approach that begins will the identification and testing of coupon-sized material and adhesive specimens. Once appropriate materials and adhesives have been verified for aerocapture in relevant environments, larger scale system-level components and finally a representative aeroshell prototype will be demonstrated. The goal will be to advance the TRL of high-temperature structural systems to a TRL of 6, and to maximize their impact on all aspects of aeroshell design.

2. NASA Ames Research Center - Development of Aeroshell Technologies for Aerocapture Missions to the Outer Planets

NASA Ames Research Center (ARC) and our team members propose to develop the critical aeroshell technologies to enable aerocapture missions in support of Solar System Exploration. The critical technologies that enable aerocapture are guidance, navigation and control (GN&C) and the thermal protection system (TPS). If the Aerocapture Flight Test Experiment (AFTE) was selected for the New Millennium Program ST7 mission (and was successful), much of the GN&C risk for aerocapture could have been retired. Unfortunately, AFTE was not selected for ST7. Independently, however, the TPS issues for outer planet exploration have not been adequately addressed and require a significant development effort. An aerocapture mission to Titan can be accomplished in the near-term utilizing existing TPS materials, some of which are flight proven. There are other materials, not yet at TRL 6, which offer potential cost savings and/or an enhanced opportunity to acquire valuable flight test data. These will require modest further development to reach TRL 6. Our proposed developmental efforts will offer a range of TPS options for a Titan aerocapture mission. An aerocapture mission to Neptune cannot be accomplished with existing TPS materials at acceptable aeroshell mass fractions. The Neptune aerocapture flight environment will expose the TPS to both very high heat fluxes and extremely large total heat loads (factors of 2 - 5 higher than Galileo). We propose to develop a new class of TPS concepts (utilizing existing materials) that will provide reliable, efficient performance for such severe environments. TPS selection, performance, and sizing for any atmospheric entry mission is governed by definition of the aerothermal environment it will be exposed to. Existing uncertainties in current capabilities to model several flowfield characteristics result in aeroshell designs incorporating significant margins. Some of these uncertainties can only be mitigated with flight data due to limitations of existing ground test facilities. We propose to improve the aerothermal environment modeling in those areas that can be supported with focused ground tests. Our primary methodology to develop the above mentioned aeroshell technologies is utilization of arc jet, ballistic range and wind tunnel testing. The technology developed will focus on the two reference missions but will have broad application for future missions.

3. ELORET - MicroSensor and Instrumentation Technology for Aerocapture

Eloret Corp. and our partners NASA Ames Research Center (ARC), JPL, and SRI International, proposes to develop instrumentation technology for risk reduction of aerocapture for Solar System Exploration. This effort falls under the critical technology of Thermal Protection Systems. Current technology of sensors applied in flight for aeroshell flowfield or TPS response is based on 1960s technology. This effort will focus on raising the TRL from 3 to 6 for two instruments with the potential to have large impact on risk reduction and science payload mass fraction: heat flux/temperature; and recession. The recession sensor is proposed as an option to modernize the recession depth sensor flown on the Galileo mission. Direct measurement of heat flux will enable the quantification of time and location of transition to turbulence: a critical parameter in the conservative sizing of aftshell TPS mass. Both measurements will allow for direct verification of physics models used for TPS sizing, and atmospheric conditions. The baseline effort is the heat flux/temperature sensor, followed in priority by modernization of the Galileo ablation sensor option.

4. Applied Research Associates - Advanced Ablator Families for Aeroassist Missions

NRA 02-OSS-1 - In-Space Propulsion Technologies (Aerocapture) Notice of Intent to Propose Engineering Optimization, Testing and Analysis of Advanced Families of Charring Ablators for Aeroassist Missions to the Outer Planets and their Moons William M. Congdon Ablatives Laboratory Applied Research Associates, Inc. 14824 E. Hinsdale Ave., Ste. C Englewood, Colorado 80112 303/699-7737 - bcongdon@msn.com Donald M. Curry Structures and Mechanics Division NASA Johnson Space Center Houston, Texas 77058 281/483-8865 - donald.m.curry1@jsc.nasa.gov Timothy J. Collins Structures and Materials Division NASA Langley Research Center Hampton, VA 23681 757/864-3113 - t.j.collins@larc.nasa.gov Missions to the outer planets and their moons involving aeroassist technologies pose stringent thermal protection system requirements for high-performance, very lightweight ablative materials. Over the past three years, ARA, Inc. has developed family systems of advanced ablators that can meet these requirements. Now at a TRL of 4 to 5, these ablator families consist of flexible, filled silicones and phenolics with densities from 11 lb/ft3 to 28 lb/ft3, and include formulations of still higher density for the most severe entry-heating environments. The new silicone-based materials are suitable for entry environments with peak heating up to about 200 Btu/ft2-sec (227 w/cm2). One advanced formulation, Hyperlite-A at 13 lb/ft3, shows good performance and the potential for significant weight savings for a Titan aerocapture or entry mission. Lightweight flexible phenolics, such as PhenCarb-28 at 28 lb/ft3, look promising for Neptune missions with peak heating in the range from about 1,000 to 2,000 Btu/ft2-sec (~1.1 to 2.3 kw/cm2). A Phencarb-32 or -36 should have more optimal performance and efficiency for still higher heating. All of these materials in the advanced ablator families use a new, state-of-the-art manufacturing process known as the Strip-Collar Bonding Approach, or "SCBA." This involves CAD/CAM and CNC laser milling to produce high-tolerance, reliable heat shields with greatly reduced labor requirements compared to the traditional method of honeycomb bonding and packing. Demonstration, subscale manufacturing to date has shown that SCBA offers cost savings in the range of 33% compared to the honeycomb method as used on Pathfinder and MER missions. Having developed these ablator materials to an intermediate TRL, and demonstrated their efficacy for advanced planetary missions, it is now important to increase their TRL to 6 by engineering optimization, testing and analysis. Thus, we have defined a joint ARA and NASA/JSC and NASA/LaRC program under ROSS-2002 to accomplish this goal. ARA will lead this two-year project and ARA's W.M.Congdon will be the Principal Investigator. In short, ARA's role will be centered on advanced TPS materials and their manufacture, and understanding how changes in formulation and constituents produce the desired changes in performance and weight. ARA will also have leadership for understanding and correlating ablator test results and for the refining of critically-important thermal-response models (math models) for predicting TPS performance in a flight environment. NASA/JSC's project responsibility will be based on its many unique capabilities for the systems and structures engineering aspects of spacecraft heat-shield development. JSC engineering will collaborate with other NASA centers (particularly LaRC) to define entry environments for key missions as they are evolved, will address aeroshell shapes and structures, and will conduct flight analyses to evaluate/define thermal protection system performance and requirements for advanced optimization of materials and models. JSC engineering will also define the material testing required for advanced ablator systems to demonstrate a technology maturity equivalent to a TRL 6 and prescribe the database needed to extend the TRL to an 8-9 value. NASA/LaRC will be responsible for investigating and defining suitable aeroshell structures and structural materials for Titan/Neptune missions. This joint ARA/JSC/LaRC project will involve a substantial amount of arc-jet testing of advanced ablative materials. The purpose of this testing is fourfold: 1) to understand how existing advanced ablators perform for defined and simulated mission environments; 2) to rank performance and select a best candidate for each specific mission; 3) to develop data that might lead to advanced optimization of a specific candidate as needed; and 4) to generate high-fidelity thermal response data for enhancement of material response models in a closely simulated environment of specific flight missions. Arc-jet testing will use the NASA/ARC facilities, in particular the 60 MW Interaction Heating Facility. Approximately 160 ablator samples will be arc-jet tested during this two-year project.

5. Lockheed Martin Astronautics - Aerocapture Technologies

In response to the NRA 02-OSS-01, Lockheed Martin with its extensive experience in design, manufacturing and test of aeroshell, is formally submitting this proposal toward the Aerocapture Technology of the NRA. Our experiences in design and manufacturing of many NASA's planetary entry vehicles, such as Mars Pathfinder, Galileo Probe and aeroshells for the Mars Exploration Program, will help advance the entry heatshield technology to highly mass-efficient systems for outer planet aerocapture. In addition to our in-house IR&D in Entry Systems, we will rely on our flight vehicle development experience to establish an efficient set of test programs, with structural and high-energy thermal testing facilities, to improve the level of aerocapture technology. These tests will include material properties definition, system mechanical definition and high-energy arcjet testing. Only by improving the mass efficiency, effective Thermal Protection Systems of the entry aeroshells, can the Titan and Neptune Aerocapture Missions be realized. This reentry technology also has direct application to other missions such as advanced Mars landers, Earth sample return missions and the like; will benefit many other NASA missions. Our proposal concentrates on the requested four technical areas associated within Section A, the Aeroshell Technologies task. These areas are collected into technical disciplines as follows: 1) Aeroshell System Requirements, 2) Aeroshell Design, 3) TPS Materials, and 4) Structural Materials and Adhesives. The first two areas provide aerocapture technical support and aeroshell design parameters that feed the last two areas, which perform development and test of candidate systems. We have identified three specific Structure/TPS systems that we believe to be highly promising for the large range of heating conditions specified in the Titan and Neptune reference aerocapture missions: 1) Graphite composite/SLA-561V ablator system, 2) Carbon-carbon (C-C) monocoque system, 3) Option 3 advances the C-C design that incorporates high efficiency thermal insulation technology.

6. Ball Aerospace and Technologies Corp. - Technology Development of Ballute Aerocapture

Our team (Ball Aerospace & Technologies Corp., Langley Research Center, ILC Dover and the Jet Propulsion Lab) proposes to advance the critical technologies of system level trailing ballute aerocapture design from TRL 3 to TRL 6. Our core team has over 48 years of experience with aeroassist mission and technology development and 13 years of involvement with inflatable lightweight film assemblies. Our team has a mature working relationship based on current ballute aerocapture work being performed under a Gossamer NRA contract. Our trailing ballute aerocapture concepts offer payload mass fractions of more than 80%, nearly a 3 to 1 improvement over chemical orbit insertion for missions such as Titan and Neptune. The investments proposed not only retire risk and maturity issues for Titan and Neptune, but offer technology advancement and verification approaches that are scalable to many other missions. Our concepts provide complimentary technology to low-thrust advanced propulsion systems such as electric and solar sail, resulting in mass savings that enables lower cost, higher science return missions to most of the solar system. We propose a combination of design, analysis, process development and test that will retire the critical risks recognized by the aerocapture community. The critical risks addressed include: hypersonic/rarefied flow stability including wake shock focusing, aeroelasticity and structural dynamics including ballute deformation/ flowfield interactions, mission design and robust trajectory modulation including ballute separation algorithm development, tether design and ballute attachment, packaging, storage, inflation and deploy dynamics and materials assessments including strength, thermal capability, density, maturity, seaming, and fabrication. Analyses proposed include CFD/DSMC, FEA, and trajectory (3-dof and 6-dof Monte Carlo). Testing will include materials properties tests, hypersonic wind tunnel and under expanded free jet testing, along with the construction of prototypes for hypersonic tests, and deployment and inflation testing. At the end of the proposed effort we will deliver an optimized ballute configuration with minimum system mass and a roadmap for the next step: a low cost Earth flight demonstration minimizing mission and developmental risk.