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NASA NTRSDate not supplied

Advanced Supersonic Parachute Inflation Research Experiment 2 (ASPIRE2) Flight Mechanics and Parachute Performance

The Advanced Supersonic Parachute Inflation Re-search Experiment-2 (ASPIRE2) program consists of a sounding rocket flight test to be launched at Wallops Flight Facility (WFF) in 2025. This effort seeks to qualify the supersonic parachute for the Sample Retrieval Lander (SRL) under the Mars Sample Return (MSR) program, building off the successes of its predecessor program, ASPIRE, which launched in both 2017 and 2018. The conops for the ASPIRE2 mission are shown in Figure 1 relative to its predecessor. While the ASPIRE pro-gram qualified a 21.5 m diameter disk-gap band parachute deployed at Mach 1.7, the SRL parachute will be increased to 24 m in diameter and target a Mach 2.1 parachute deployment condition, the largest diameter supersonic parachute deployed at the highest Mach number to date. These requirements are driven by the increased landing mass of the sample retrieval lander, which is roughly a 50% increase over that of the Mars2020 entry vehicle. These changes necessitate ASPIRE2 to certify the parachute performance under similar deployment conditions. This work will cover the 6-DoF para-chute model implementation and flight mechanics performance for the ASPIRE2 campaign and discuss how these results will impact the design of the test article. The ASPIRE2 trajectory is simulated using a multi-body flight dynamics tool, Program to Optimize Simulated Trajectories II (POST2). This work will summarize both the flight mechanics and parachute modeling (from payload separation until splashdown) and performance for this mission. Various design trades will be assessed such as para-chute tuning parameters as well as vehicle center of gravity location.

Evan Roelke↗

NASA NTRS2016

Conventional and Bimodal Nuclear Thermal Rocket (NTR) Artificial Gravity Mars Transfer Vehicle Concepts

A variety of countermeasures have been developed to address the debilitating physiological effects of zero-gravity (0-g) experienced by cosmonauts and astronauts during their approximately 0.5 to 1.2 year long stays in low Earth orbit (LEO). Longer interplanetary flights, combined with possible prolonged stays in Mars orbit, could subject crewmembers to up to approximately 2.5 years of weightlessness. In view of known and recently diagnosed problems associated with 0-g, an artificial gravity (AG) spacecraft offers many advantages and may indeed be an enabling technology for human flights to Mars. A number of important human factors must be taken into account in selecting the rotation radius, rotation rate, and orientation of the habitation module or modules. These factors include the gravity gradient effect, radial and tangential Coriolis forces, along with cross-coupled acceleration effects. Artificial gravity Mars transfer vehicle (MTV) concepts are presented that utilize both conventional NTR, as well as, enhanced bimodal nuclear thermal rocket (BNTR) propulsion. The NTR is a proven technology that generates high thrust and has a specific impulse (Isp) capability of approximately 900 s-twice that of today's best chemical rockets. The AG/MTV concepts using conventional Nuclear Thermal Propulsion (NTP) carry twin cylindrical International Space Station (ISS)- type habitation modules with their long axes oriented either perpendicular or parallel to the longitudinal spin axis of the MTV and utilize photovoltaic arrays (PVAs) for spacecraft power. The twin habitat modules are connected to a central operations hub located at the front of the MTV via two pressurized tunnels that provide the rotation radius for the habitat modules. For the BNTR AG/MTV option, each engine has its own closed secondary helium(He)-xenon (Xe) gas loop and Brayton Rotating Unit (BRU) that can generate 10s of kilowatts (kWe) of spacecraft electrical power during the mission coast phase eliminating the need for large PVAs. A single inflatable TransHab-type habitation module is also used with multiple vertical floors oriented radial to the MTV spin axis. The BNTR MTV's geometry-long and linear-is naturally compatible with AG operation. By rotating the vehicle about its center-of-mass (CM) and perpendicular to its flight vector at approximately 3.0 to 5.2 rpm, a centrifugal force and AG environment corresponding to approximately 0.38 to 1.0 g can be established to help maintain crew fitness out to Mars and back. Vehicles using NTP/ Bimodal Nuclear Thermal Propulsion (BNTP) can more readily accommodate the heavier payload mass and increased RCS propellant loading associated with AG operation, and can travel faster to and from Mars thereby reducing the crew's exposure to galactic cosmic radiation and solar flares. Mission scenario descriptions, key vehicle features and operational characteristics for each propulsion option are presented using the lift capability and payload volumes estimated for the Space Launch System (SLS)-1B and followon Heavy Lift Vehicle (HLV).

Nuclear Propulsion↗

NASA NTRS2012

Modified Fittings Enhance Industrial Safety

Kennedy Space Center is not only home to one of the largest buildings in the world - the massive Vehicle Assembly Building - it also hosts a number of one-of-a-kind facilities. The more than 30-mile-long campus has witnessed every launch from the Space Shuttle Launch Pad, as well as many homecomings at the Shuttle Landing Facility. Just as important, the Space Station Processing Facility (SSPF) has seen each element of the International Space Station (ISS) that passes through Kennedy before it goes into orbit. The SSPF is where ISS components are checked, tested, and adjusted before being packed into the Space Shuttle for transport. In an environment like the SSPF - spanning 457,000 square feet of processing areas, operational control rooms, laboratories, logistics areas, and office space - large workstands and equipment used to support the processing of ISS components need to be moved around the facility. One of the devices employed for this task is an air pallet. An air pallet moves on cushions of air instead of wheels. Compressed air inflates the cushions underneath the pallet and is then expelled through exhaust holes. This forms a thin film of air between the cushions and the floor, lifting the platform off the floor and making it easy to move the heavy workstands, equipment, and ISS components. Concerned with the safety of the connections on the pressurized air hoses used for the air pallets, engineers at Kennedy modified an existing commercial cam and groove fitting to control the air supply hose in the event of an accidental release of a pressurized hose. This modification prevented the hose from detaching and, propelled by compressed air, striking workers or equipment. "At the time, these were not available on commercial coupling halves, so NASA made a modification and then put them into use. If a worker were to accidentally try to remove a pressurized hose from the pallet, it no longer rapidly separated, and it safely relieved the pressure," says Paul Schwindt, an engineer at Kennedy who together with Alan Littlefield, also an engineer at Kennedy, designed the modification.

Source record↗

NASA NTRS2010

Gossamer Technology to Deorbit LEO Non-Propulsion Fitted Satellite

Since 2004, CNES has decided to apply the end of life Code of Conduct rules to debris mitigation. Originally drawn up by the main European space agencies, it contains basic rules to be applied in space in order to limit the increase of orbital debris. In low Earth orbit, the rule is to limit in-orbit lifetime to 25 years after the end of the operational mission, or else to transfer to a graveyard orbit above 2000 km. In order to follow these instructions, a task force was set up in 2005 to find the best way to implement them on MICROSCOPE and CNES microsatellite family (MYRIADE). This 200-kg spacecraft should be launched in 2014 on a 790-km high circular orbit. Without targeted action, its natural re-entry would occur in 67 years. Two strategies to reduce this time period were compared: propulsive maneuvers at the end of the mission or the deployment of large surfaces to increase significantly the ballistic coefficient. At the end of the trade off, it was recommended: .. For the non-propulsive system fitted satellites, to use passive aerobraking by deployment of added surface, .. For satellites having propulsive subsystem in baseline for mission purposes, to keep sufficient propellant and implement specific maneuvers. The poster gives an overview of the process that led to the development of a deployable aerobraking wing using a lightweight aluminized Kapton membrane and an inflatable aluminum laminate boom. The main requirements; The trade off among various aerobraking solutions; The development plan. This technology presents a very attractive potential and it could be a first step in using of inflatable technology on spaces vehicles, before to deal with others more exigent applications.

Dupuy, C.↗

NASA NTRS1998

Contamination Control Considerations for the Next Generation Space Telescope (NGST)

The NASA Space Science Program, in its ongoing mission to study the universe, has begun planning for a telescope that will carry on the Hubble Space Telescope's exploration. This telescope, the 'Next Generation Space Telescope' (NGST), will be 6-8 meters in diameter, will be radiatively cooled to 30-60 Kelvin in order to enable extremely deep exposures at near infrared wavelengths, and will operate for a lifetime of 5-10 years. The requirement will be to measure wavelengths from 1-5 microns, with a goal to measure wavelengths from 0.6-30 microns. As such, NGST will present a new contamination control challenge. The Goddard Space Flight Center (GSFC) performed one of three preliminary feasibility studies for the NGST, presenting a telescope with an 8 meter, deployable primary mirror and a deployable secondary mirror. The telescope would be radiatively cooled, with the optical telescope assembly (OTA) and the science instrument module (SIM) isolated from the warmer spacecraft support module (SSM). The OTA and the SIM would also be shielded from sunlight with an enormous, inflatable sun-shield. The GSFC telescope was designed for launch on an Atlas HAS, which would require launching the telescope in a stowed configuration, with the SSM, antennae, sun-shield, primary mirror 'petals', and secondary mirror deployed once on-orbit. The launch configuration and deployment scenario of an exposed telescope measuring near infrared and cooled to 30-60 K are the factors presenting contamination hazards to the NGST mission. Preliminary science requirements established are: less than 20% reflectance decrease on optical surfaces over the wavelength range, and less than 0.3% obscuration of optical surfaces. In order to meet these requirements, NGST must be built and launched with careful attention to contamination control. Initial contamination control design options include strict selecting of materials and baking out of hardware down to the component level, minimizing or eliminating exposure of the OTA to sunlight or earth albedo during deployment and early on-orbit operations, cleaning of the primary and secondary mirrors at the launch site, cleaning of the launch vehicle fairing, locating thrusters and vents on the warm side of the sun shield only, and the possibility of including a deployable cover if that is shown to be necessary.

Wooldridge, Eve M.↗

NASA NTRS2006

Aerocapture Benefits to Future Science Missions

NASA's In-Space Propulsion Technology (ISPT) Program is investing in technologies to revolutionize the robotic exploration of deep space. One of these technologies is Aerocapture, the most promising of the "aeroassist" techniques used to maneuver a space vehicle within an atmosphere, using aerodynamic forces in lieu of propellant. (Other aeroassist techniques include aeroentry and aerobraking.) Aerocapture relies on drag atmospheric drag to decelerate an incoming spacecraft and capture it into orbit. This technique is very attractive since it permits spacecraft to be launched from Earth at higher velocities, providing shorter trip times and saving mass and overall cost on future missions. Recent aerocapture systems analysis studies quantify the benefits of aerocapture to future exploration. The 2002 Titan aerocapture study showed that using aerocapture at Titan instead of conventional propulsive capture results in over twice as much payload delivered to Titan. Aerocapture at Venus results in almost twice the payload delivered to Venus as with aerobraking, and over six times more mass delivered into orbit than all-propulsive capture. Aerocapture at Mars shows significant benefits as the payload sizes increase and as missions become more complex. Recent Neptune aerocapture studies show that aerocapture opens up entirely new classes of missions at Neptune. Current aerocapture technology development is advancing the maturity of each subsystem technology needed for successful implementation of aerocapture on future missions. Recent development has focused on both rigid aeroshell and inflatable aerocapture systems. Rigid aeroshell systems development includes new ablative and non-ablative thermal protection systems, advanced aeroshell performance sensors, lightweight structures and higher temperature adhesives. Inflatable systems such as trailing tethered and clamped "ballutes" and inflatable aeroshells are also under development. Computational tools required to support future aerocapture missions are an integral part of aerocapture development. Tools include engineering reference atmosphere models, guidance and navigation algorithms, aerothermodynamic modeling, and flight simulation.

Artis, Gwen↗

NASA NTRS2020

Parametric Cost Estimates of Four 20 Ton Payload Mars EDL Vehicle Concepts

Cost is one of the biggest obstacles to sending humans to Mars. Many studies have examined various vehicles and/or technologies which could be used to send humans to Mars. However, nearly all of these studies have either ignored cost or have used the mass of the vehicles as a surrogate for cost. This work presents parametric cost estimates of four Mars Entry, Descent, and Landing vehicle concepts capable of landing 20 tons of payload on the surface of Mars. The concepts are the Co-Optimization Blunt-body Re-entry Analysis-Mid lift to drag ratio Rigid Vehicle (Cobra–MRV), Adaptive Deployable Entry and Placement Technology (ADEPT), Hypersonic Inflatable Aerodynamic Decelerator (HIAD), and Capsule concepts. Two variants of the HIAD concept are studied, one with a single HIAD heatshield and one with two HIADs. The study presents the development and production cost of each vehicle and the total cost to build five flight units and fifteen flight units. Five flight units is representative of a single mission, or a "Flags and Footprints" style campaign, whereas fifteen flight units is representative of multiple missions going to multiple sites on Mars. All costs presented in this work are normalized to the average of the first unit production cost of the lander concepts. There is a large amount of uncertainty both in the design of the vehicles and the cost of the vehicles and it is impossible to claim with certainty which vehicle will cost the least. However, it is likely that the total cost to produce five flight units will be in the following ranking from least costly to most costly, Single HIAD, Capsule, Dual HIAD, Cobra, ADEPT. The ranking for fifteen units is similar but with the positions of the Dual HIAD and Cobra reversed. This study only considers the cost of the entry vehicle and does not take into account the effect of the entry vehicle’s mass on the cost of the in-space propulsion stage or launch vehicle.

Paul D Friz↗

NASA NTRS2009

Findings from the Supersonic Qualification Program of the Mars Science Laboratory Parachute System

In 2012, the Mars Science Laboratory Mission (MSL) will deploy NASA's largest extra-terrestrial parachute, a technology integral to the safe landing of its advanced robotic explorer on the surface. The supersonic parachute system is a mortar deployed 21.5 m disk-gap-band (DGB) parachute, identical in geometric scaling to the Viking era DGB parachutes of the 1970's. The MSL parachute deployment conditions are Mach 2.3 at a dynamic pressure of 750 Pa. The Viking Balloon Launched Decelerator Test (BLDT) successfully demonstrated a maximum of 700 Pa at Mach 2.2 for a 16.1 m DGB parachute in its AV4 flight. All previous Mars deployments have derived their supersonic qualification from the Viking BLDT test series, preventing the need for full scale high altitude supersonic testing. The qualification programs for Mars Pathfinder, Mars Exploration Rover, and Phoenix Scout Missions were all limited to subsonic structural qualification, with supersonic performance and survivability bounded by the BLDT qualification. The MSL parachute, at the edge of the supersonic heritage deployment space and 33% larger than the Viking parachute, accepts a certain degree of risk without addressing the supersonic environment in which it will deploy. In addition, MSL will spend up to 10 seconds above Mach 1.5, an aerodynamic regime that is associated with a known parachute instability characterized by significant canopy projected area fluctuation and dynamic drag variation. This aerodynamic instability, referred to as "area oscillations" by the parachute community has drag performance, inflation stability, and structural implications, introducing risk to mission success if not quantified for the MSL parachute system. To minimize this risk and as an alternative to a prohibitively expensive high altitude test program, a multi-phase qualification program using computation simulation validated by subscale test was developed and implemented for MSL. The first phase consisted of 2% of fullscale supersonic wind tunnel testing of a rigid DGB parachute with entry-vehicle to validate two high fidelity computational fluid dynamics (CFD) tools. The computer codes utilized Large Eddy Simulation and Detached Eddy Simulation numerical approaches to accurately capture the turbulent wake of the entry vehicle and its coupling to the parachute bow-shock. The second phase was the development of fluid structure interaction (FSI) computational tools to predict parachute response to the supersonic flow field. The FSI development included the integration of the CFD from the first phase with a finite element structural model of the parachute membrane and cable elements. In this phase, a 4% of full-scale supersonic flexible parachute test program was conducted to provide validation data to the FSI code and an empirical dataset of the MSL parachute in a flight-like environment. The final phase is FSI simulations of the full-scale MSL parachute in a Mars type deployment. Findings from this program will be presented in terms of code development and validation, empirical findings from the supersonic testing, and drag performance during supersonic operation.

EDL↗

NASA NTRS2004

Design and Experimental Verification of Deployable/Inflatable Ultra-Lightweight Structures

Because launch cost of a space structural system is often proportional to the launch volume and mass and there is no significant gravity in space, NASA's space exploration programs and various science missions have stimulated extensive use of ultra-lightweight deployable/inflatable structures. These structures are named here as Highly Flexible Structures (HFSs) because they are designed to undergo large displacements, rotations, and/or buckling without plastic deformation under normal operation conditions. Except recent applications to space structural systems, HFSs have been used in many mechanical systems, civil structures, aerospace vehicles, home appliances, and medical devices to satisfy space limitations, provide special mechanisms, and/or reduce structural weight. The extensive use of HFSs in today's structural engineering reveals the need of a design and analysis software and a database system with design guidelines for practicing engineers to perform computer-aided design and rapid prototyping of HFSs. Also to prepare engineering students for future structural engineering requires a new and easy-to- understand method of presenting the complex mathematics of the modeling and analysis of HFSs. However, because of the high flexibility of HFSs, many unique challenging problems in the modeling, design and analysis of HFSs need to be studied. The current state of research on HFSs needs advances in the following areas: (1) modeling of large rotations using appropriate strain measures, (2) modeling of cross-section warpings of structures, (3) how to account for both large rotations and cross- section warpings in 2D (two-dimensional) and 1D structural theories, (4) modeling of thickness thinning of membranes due to inflation pressure, pretension, and temperature change, (5) prediction of inflated shapes and wrinkles of inflatable structures, (6) development of efficient numerical methods for nonlinear static and dynamic analyses, and (7) filling the gap between geometrically exact elastic analysis and elastoplastic analysis. The objectives of this research project were: (1) to study the modeling, design, and analysis of deployable/inflatable ultra-lightweight structures, (2) to perform numerical and experimental studies on the static and dynamic characteristics and deployability of HFSs, (3) to derive guidelines for designing HFSs, (4) to develop a MATLAB toolbox for the design, analysis, and dynamic animation of HFSs, and (5) to perform experiments and establish an adequate database of post-buckling characteristics of HFSs.

Pai, P. Frank↗

NASA NTRSDate not supplied

Capability Gaps Assessment and Identification of Critical Technology Elements for Mars Transit Habitat

The Habitation Systems Development Office (HP40) at NASA Marshall Space Flight Center supports systems engineering, integration, and project management for next generation space habitats. For in space operations and eventual transport of humans to Mars, NASA will rely on a Mars Transit Habitat (TH). The TH will be designed for an up to 1,200-day Mars mission and will carry all food and supplies needed to support four crew for this duration. In the current concept of operations, Mars TH transfers to near rectilinear halo orbit (NRHO) following launch and docks at Gateway as a visiting vehicle. While there, the TH will complete system shakedown testing and a series of analog missions which will grow from 3 to 6+ months in duration TH also augments Gateway’s habitation capabilities beyond 60-days. Proposed Gateway-TH missions will far exceed the longest duration cislunar human missions to date. These shakedown missions will also be the first operational readiness tests of Mars TH’s long-duration deep space systems, and of the split crew (two crew on the surface, two crew in space) operations that are vital to the approach for the first human Mars mission. Once shakedown missions are complete, Mars TH departs Gateway to aggregate with the Mars propulsion system in NRHO before onboarding the crew and final supplies in Earth orbit via a co-manifested Orion-logistics module. Orion and the LM return to Earth prior to the now aggregated Deep Space Transport vehicle’s journey to Mars. Development of the Mars TH requires significant technology development and maturation. Each year the agency performs a capability gaps assessment, where gaps developed by subject matter experts (SMEs) in various engineering/science disciplines are linked to architectural elements in formulation and prioritized. A gap captures the difference between the current state-of-the-art and the maturity of the capability that is needed to enable or enhance a mission as it is currently envisioned in the government reference architecture. HP40 conducted a gap analysis for Mars TH which will be summarized in this poster. Gaps classified as enabling (which means the mission cannot achieve success without gap closure) were subsequently used to identify critical technology elements (CTEs) for Mars TH. This identification of CTEs was also informed by an examination of the product breakdown structure for Mars TH and focused conversations with SMEs in specific technology areas. CTEs identified for Mars TH to date include the following (note this is not a comprehensive list – CTEs listed represent those in MSFC’s capability areas): inflatable softgoods for habitation; enhanced CO2 recovery; life support systems with greater levels of reliability and maintainability; autonomous guidance, navigation, command and control; and radiators for the Mars TH application. The habitation systems development team is currently delving deeper into each CTE to assess technology approaches being pursued, their maturity, and the degree of difficulty in maturation to meet projected Mars TH timelines. This poster will summarize work to date on the identification of enabling capability gaps linked to Mars TH and provide insight into the associated CTEs and technology maturation efforts.

technology development↗

NASA NTRS2013

Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology Development Overview

The successful flight of the Inflatable Reentry Vehicle Experiment (IRVE)-3 has further demonstrated the potential value of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology. This technology development effort is funded by NASA's Space Technology Mission Directorate (STMD) Game Changing Development Program (GCDP). This paper provides an overview of a multi-year HIAD technology development effort, detailing the projects completed to date and the additional testing planned for the future. The effort was divided into three areas: Flexible Systems Development (FSD), Mission Advanced Entry Concepts (AEC), and Flight Validation. FSD consists of a Flexible Thermal Protection Systems (FTPS) element, which is investigating high temperature materials, coatings, and additives for use in the bladder, insulator, and heat shield layers; and an Inflatable Structures (IS) element which includes manufacture and testing (laboratory and wind tunnel) of inflatable structures and their associated structural elements. AEC consists of the Mission Applications element developing concepts (including payload interfaces) for missions at multiple destinations for the purpose of demonstrating the benefits and need for the HIAD technology as well as the Next Generation Subsystems element. Ground test development has been pursued in parallel with the Flight Validation IRVE-3 flight test. A larger scale (6m diameter) HIAD inflatable structure was constructed and aerodynamically tested in the National Full-scale Aerodynamics Complex (NFAC) 40ft by 80ft test section along with a duplicate of the IRVE-3 3m article. Both the 6m and 3m articles were tested with instrumented aerodynamic covers which incorporated an array of pressure taps to capture surface pressure distribution to validate Computational Fluid Dynamics (CFD) model predictions of surface pressure distribution. The 3m article also had a duplicate IRVE-3 Thermal Protection System (TPS) to test in addition to testing with the Aerocover configuration. Both the Aerocovers and the TPS were populated with high contrast targets so that photogrammetric solutions of the loaded surface could be created. These solutions both refined the aerodynamic shape for CFD modeling and provided a deformed shape to validate structural Finite Element Analysis (FEA) models. Extensive aerothermal testing has been performed on the TPS candidates. This testing has been conducted in several facilities across the country. The majority of the testing has been conducted in the Boeing Large Core Arc Tunnel (LCAT). HIAD is continuing to mature testing methodology in this facility and is developing new test sample fixtures and control methodologies to improve understanding and quality of the environments to which the samples are subjected. Additional testing has been and continues to be performed in the NASA LaRC 8ft High Temperature Tunnel, where samples up to 2ft by 2ft are being tested over representative underlying structures incorporating construction features such as sewn seams and through-thickness quilting. With the successful completion to the IRVE-3 flight demonstration, mission planning efforts are ramping up on the development of the HIAD Earth Atmospheric Reenty Test (HEART) which will demonstrate a relevant scale vehicle in relevant environments via a large-scale aeroshell (approximately 8.5m) entering at orbital velocity (approximately 7km/sec) with an entry mass on the order of 4MT. Also, the Build to Print (BTP) hardware built as a risk mitigation for the IRVE-3 project to have a "spare" ready to go in the event of a launch vehicle delivery failure is now available for an additional sub-orbital flight experiment. Mission planning is underway to define a mission that can utilize this existing hardware and help the HIAD project further mature this technology.

Hughes, Stephen J.↗

NASA NTRS2001

SHARP-B2: Flight Test Objectives, Project Implementation and Initial Results

On September 28, 2000 the SHARP-B2 flight experiment was launched from Vandenberg Air Force Base, California. SHARP-B2 is the 2nd Ballistic flight test in the SHARP (Slender Hypervelocity Aerothermodynamic Research Probes) program which develops and tests new thermal protection materials and sharp body concepts. The flight tested Ultra-High Temperature Ceramics (UHTCs), which may radically change the design and performance of future aerospace vehicles. The new designs may overturn an age-old tenet of aerodynamics: that blunt-body aerospace vehicles, but not those with sharp leading edges, can survive the searing temperatures created as the vehicles tear through the atmosphere. Sharp leading edges offer numerous advantages over the blunt-body design currently in use. They could allow a space shuttle or crew return vehicle to maneuver in space more like an airplane and potentially allow astronauts to return to Earth from anywhere on orbit. They may allow improved astronaut safety by decreasing the risk of aborting into the ocean. They may reduce the electromagnetic interference that causes the communications blackouts that plague reentering blunt-body space vehicles. Reducing the amount of drag could lead to a reduction in propulsion requirements. Planetary probes could make use of sharp-body technology for aerobraking and to maximize their maneuvering capability. SHARP-B2 was a joint effort among NASA Ames, Sandia National Laboratories, the U.S. Air Force and the U.S. Army. It was funded by the Pathfinder Program at NASA's Marshall Space Flight Center. The SHARP-B2 payload was carried aboard a U.S. Air Force Minuteman III missile carrying a modified Mk 12A reentry vehicle (RV), which blasted off from Vandenberg Air Force Base near Lompoc, CA, at 3:01 a.m. PDT on Sept. 28. The RV was equipped with four 5. 1 inch-long strakes, or sharp leading edges. Each strake contained three UHTCs: ZrB2/SiC/C; ZrB2/SiC; and HfB2/SiC. Once it reached an altitude of about 400 nautical miles, the RV was released, returning through Earth's atmosphere at speeds exceeding Mach 22. One pair of strakes was designed to retract just before reaching temperatures high enough to cause the material to begin ablating. The other pair was designed to retract shortly after ablation began, at an expected temperature of nearly 5,100 degrees Fahrenheit. Sensors in the strakes measured how closely performance matched pre-flight calculations, and data was successfully collected throughout the 23-minute flight. A parachute was deployed (but not fully inflated) and the RV splashed down in a lagoon at the Kwajalein missile range in the Pacific Ocean. Within 3 hours radar track analysis showed ocean entry to be precisely at the latitude/longitude coordinates estimated during pre-flight simulation. An hour later a ship was deployed by the Army to recover the reentry vehicle, which was recovered in 165 feet of water, just 500 feet from its planned splash-down point. This is the first RV recovery in over a decade.

Salute, Joan↗

NASA NTRS2014

Nuclear Thermal Rocket (NTR) Development Risk Communication

There are clear advantages of development of a Nuclear Thermal Rocket (NTR) for a crewed mission to Mars. NTR for in-space propulsion enables more ambitious space missions by providing high thrust at high specific impulse (approximately 900 sec) that is 2 times the best theoretical performance possible for chemical rockets. Missions can be optimized for maximum payload capability to take more payload with reduced total mass to orbit; saving cost on reduction of the number of launch vehicles needed. Or missions can be optimized to minimize trip time significantly to reduce the deep space radiation exposure to the crew. NTR propulsion technology is a game changer for space exploration. However, "NUCLEAR" is a word that is feared and vilified by some groups and the hostility towards development of any nuclear systems can meet great opposition by the public as well as from national leaders and people in authority. Communication of nuclear safety will be critical to the success of the development of the NTR. Why is there a fear of nuclear? A bomb that can level a city is a scary weapon. The first and only times the Nuclear Bomb was used in a war was on Hiroshima and Nagasaki during World War 2. The "Little Boy" atomic bomb was dropped on Hiroshima on August 6, 1945 and the "Fat Man" on Nagasaki 3 days later on August 9th. Within the first 4 months of bombings, 90- 166 thousand people died in Hiroshima and 60-80 thousand died in Nagasaki. It is important to note for comparison that over 500 thousand people died and 5 million made homeless due to strategic bombing (approximately 150 thousand tons) of Japanese cities and war assets with conventional non-nuclear weapons between 1942- 1945. A major bombing campaign of "firebombing" of Tokyo called "Operation Meetinghouse" on March 9 and 10 consisting of 334 B-29's dropped approximately1,700 tons of bombs around 16 square mile area and over 100 thousand people have been estimated to have died. The declaration of death is very clear for conventional weapons and then the declaration of death due to radiation becomes vague and unclear. This may have been due to people mis-understanding the dangers and effects of radiation when assessing the damage and harm to people initially, but it is also become insidious when expressing opposition to nuclear energy. A nuclear radiation accident can be scary due to the power involved and the fear of radiation release. The International Atomic Energy Agency defines a nuclear and radiation accident a "an event that has led to significant consequences to people, the environment or the facility." There have been 3 commercial nuclear reactor accidents (Chernobyl, Three Mile Island, and Fukushima) that stand out to the public and much of the information about the result and impact to workers, environment, and public can be misleading. Often information is presented without clear correlation with radiation and other pertinent information is left out presenting a very scary situation to affect the emotions of the reader. A very boring but "critically acclaimed" movie was made in 1979 called "The China Syndrome" starring Jane Fonda and Jack Lemmon. The film was released on March 16, 1979, 12 days before the 3-Mile Island nuclear accident in Pennsylvania. The basis from the movie was from a few nuclear plant incidents and in particular, the Brown's Ferry Alabama Power Plant fire. In one scene from the movie, a physicist Dr. Elliott Lowell played by Donald Hotton states that a China Syndrome event would make "an area the size of Pennsylvania" permanently uninhabitable. Real serious nuclear incidents like Chernobyl and Fukushima are often sited to make people fear the consequences of using nuclear power. However, the consequences are at best poorly communicated and at worst fictitiously inflated to instigate social unrest against nuclear power. There is an article being circulated on Facebook with a title "28 Signs that the (US) West Coast is being absolutely fried with nuclear radiation from Fukushima" which focus on mis-information and fear mongering. Nuclear power and NTR are powerful resources that can open many doors for future prosperity and capability. With great power comes great responsibility. Radiation and its effects need to be better understood, quantified, and communicated. A human mission to mars has its own risks of deep space radiation and is considered a considerable risk at 400 milli-Sieverts per year in deep space and 245 milli-Sieverts per year on the surface of Mars as measured by the Mars Curiosity mission. Although these quantities of ionizing radiation are within the astronaut career limit, it exceeds the yearly average amounts of ionizing radiation. Astronaut crews have experienced these levels of radiation before, but for durations shorter than a year, and a mission to Mars could possibly be 3 years in length. There is also evidence that people can comfortably handle higher levels of ionizing radiation where the radiation occurs naturally like Ramsar, Iran when people can experience 270 milli-Sieverts per year. A risk posture that the development, test, and flight of an NTR will meet opposition from groups who oppose nuclear energy must be likely and the impact can be sever to the effort. Active risk mitigation must be taken for an NTR full-scale development project. The NTR design must take into account safety for transport and off nominal conditions. Nuclear fuel element must consider containment of fission products and Low Enriched Uranium (LEU) that may meet less opposition should be considered for safety and security reasons. Even though testing was conducted on Rover/NERVA safely and successfully in the 60's with exhaust sent heavenward in to open air, modern testing of NTR must consider full containment and no release of ionizing radiation to the public and must meet the current requirement of no more than 0.1 milli-Sieverts per year to the public. 0.1 milli-Sieverts is equivalent to eating one banana or a 20 hour plane flight. Good communication with the public and regulatory agencies will be essential to show that all effort is applied toward protection to the public and astronauts. The inspiring endeavor to put humans on Mars to study the planet, search for life, and learn more about this Solar System will be full of risks but it will be worth it. NTR will be worth the development effort if it allows humans to explore our Solar System.

Kim, Tony↗