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NASA NTRS2018

A Plasma Aerocapture and Entry System for Manned Missions and Planetary Deep Space Orbiters

A rapid mission to Mars requires a large change in vehicle velocity upon arrival to establish a stable orbit. This demand is even greater for a Neptune science, requiring many kilometers per second of V. It is clear from past mission studies that a manned Mars mission and deep space planetary orbiters require aerobraking and aerocapture which use aerodynamic drag forces to slow the spacecraft. Aerocapture would enable long term studies of the outer planets and moons that would not be possible with existing braking methodologies. While the ability to utilize these atmospheres to slow down and capture spacecraft would dramatically reduce the cost, launch mass, and travel time, currently planned approaches require significant additional spacecraft mass and risk as the spacecraft must descend deep into the planetary atmosphere in order to produce significant drag on a relatively small aeroshell. The plasma based Magnetoshell being developed in this program holds the potential to perform the desired braking with significantly increased drag and control while dramatically reducing mass. Most importantly, this technology significantly lowers the risk involved with aerocapture thereby making manned planetary missions possible. The fundamental physics of the Magnetoshell is based on demonstrated experimental results. Successful implementation will dramatically decrease radiation exposures, mission risk, launch cost, and launch mass. Implementation of aerobraking by employing a solid deflector or aeroshell as a method for orbit insertion and circularization has been successfully demonstrated in the past, resulting in launch mass savings greater than 50%. In order to reduce the effect of frictional heating and dynamic pressure on the typically fragile aeroshell, or worse solar panels, the braking must be distributed over many orbital passes at a high altitude in the less dense regions of the atmosphere. It can thus take several months for a meter-scale, 1000 kg, aeroshell to execute the many elliptic orbital passes through the atmosphere to achieve the required V. This rather slow method of braking not only reduces frictional heating and dynamic forces, but also avoids unpredictable dynamic behavior due to turbulence, as well as unknown and seasonally variable atmospheric composition and temperature which has led to dangerous, mission-critical events. For exploration class missions such as DRA 5.0 aerocapture and Thermal Protection Systems (TPS) are proposed for breaking at Mars for cargo missions. Traditional aerocapture is considered too risky for manned missions. Even with the enormous mass savings that aerocapture allows, it still requires 80 tons of aeroshell and significantly increased launch mass and propellant. As will be shown, by using Magnetoshells for aerobraking, the DRA 5.0 mission will save 224 metric tons (MT) and greater than $2 B in launch costs. Beyond the dramatic savings for existing mission architectures, a low-mass, risk-free aerocapture system would allow much more rapid missions to Mars and deep space orbiters by allowing direct, faster trajectories. As will be shown, the plasma Magnetoshell Aerobraking, Aerocapture, and Entry System (AAES) not only reduces mass and cost while enabling significant new mission architectures, but also significantly reduces radiation exposures by decreasing trip times.

Manned Missions↗
NASA NTRS1994

Ways to Increase Launch Velocities of 2-Stage Gas Guns

The amount of space debris is rapidly increasing and the debris is distributed over a wide variety of orbits. Satellites, manned space vehicles and space stations will have to pay increasing attention to the dangers of impacts with space debris. Various armoring techniques (i.e., double or triple layer armor) will have to tested extensively to determine the most effective armor per unit weight. Intersecting near-earth orbits can lead to impact velocities up to 15 km/sec. Conventional two-stage light gas guns can launch intact, controlled-shape projectiles with a density of 1.2 gm/cc and length- to-diameter ratios of 0.5-1.0 at velocities up to 8-9 km/sec. Higher velocities (10-11 km/sec) can be obtained' for very light projectiles. The higher launch velocities tend to be very severe on the high pressure coupling and barrel of the gun and lead to short component lifetimes. Clearly, the ability to raise the launch velocity of a gun (for reasonably massive projectile shapes) from 8-9 km/sec to 11-13 km/sec (or higher), without reduction of component lifetimes, would have significant benefits. This would allow much better simulation of the higher velocity debris impacts as well as better simulation of high speed re-entry into planetary atmospheres. Several techniques for increasing the launcher muzzle velocity above 8-9 km/sec have been studied using CFD simulations and appear to offer the potential for significant gains. The first technique is to use multiple compressions, instead of a single compression, in the pump tube of the light gas gun. In a sense, this is a kind of pre-heating of the gas in the pump tube; other types of pre-heating have yielded disappointing results in the past. The dynamics of the multiple compression pump tube is very different, however, from the earlier techniques, where the pump tube was typically heated ohmically before the gun 2 cycle was started. In this paper, we present CFD calculations that show that significant increases in muzzle velocity can be obtained with multiple compressions in the pump tube. With a conventional two-stage gun, an important limitation to obtaining higher velocities is friction and heat transfer to the barrel, which typically has a length- to- diameter ratio of 200-400. These viscous losses greatly reduce the effectiveness of the regions of the barrel far removed from the second stage breech. We have studied computationally the effect of adding an additional breech (or breeches) along the barrel to reduce these viscous losses. Velocity increases from 6.5 to 7.2 km/sec have been obtained using the main breech and one additional breech. In these results, both breeches were operated with hydrogen, heated electrothermally. We have also studied a gun geometry where the main breech is operated in the conventional manner, using piston compression. The additional breech is operated either with electrothermal heating or heating by using a high explosive charge in a novel geometry. The latter option provides very effective compression, heating and acceleration of the hydrogen working gas and is fully reusable. Calculations are presented which show that very substantial increases in muzzle velocity can be obtained this way, without overstressing the projectile or the 'gun. The third technique studied is to add a section of ram accelerator tube after the barrel to further accelerate the projectile. The ram accelerator used here is not the conventional premixed gas ram accelerator, but a new technique using high explosive as the energy source and pure hydrogen as the working gas in a geometry which can be made fully reusable. Preliminary results with this new rain accelerator geometry were presented and showed that stable ram accelerator drive can be established. Herein, detailed calculations axe presented which show that substantial velocity increases can be obtained using this ram accelerator technique in tandem with a conventional light gas gun.

Bogdanoff, David W.↗
NASA NTRS1976

Accretion of the terrestrial planets. II

The theory of gravitational accretion of the terrestrial planets is examined. The concept of a 'closed feeding zone' is somewhat unrealistic, but provides a lower bound on the accretion time. A velocity relation for planetesimals which includes an initial velocity component is suggested. The orbital parameters of the planetesimals and the dimensions of the feeding zone are related to their relative velocities. The assumption of an initial velocity does not seriously change the accretion time. Mercury, Venus, and the earth have accretion times on the order of 100 million years. Mars requires well over one billion years to accrete by the same assumptions. The lunar cratering history makes a late formation of Mars unlikely. If Mars is as old as the earth, nongravitational forces or a violation of the feeding zone concept is required. One such possibility is the removal of matter from the zone of Mars by Jupiter's influence. The final sweeping up by Mars would result in the scattering of a considerable mass among the other terrestrial planets. The late postaccretional bombardments inferred for the moon and Mercury may have had this source.

Weidenschilling, S. J.↗
NASA NTRS2014

Analytic Ballistic Performance Model of Whipple Shields

The dual-wall Whipple shield is the shield of choice for lightweight, long-duration flight. The shield uses an initial sacrificial wall to initiate fragmentation and melt an impacting threat that expands over a void before hitting a subsequent shield wall of a critical component. The key parameters to this type of shield are the rear wall and its mass which stops the debris, as well as the minimum pressure generated under threat particle impact of the sacrificial wall and the amount of void that is available for expansion. Ensuring the minimum pressure is sufficiently high to achieve large scale fragmentation/melt of the threat particle enables the expansion of the threat and reduces the momentum flux of the debris on the rear wall. Three key factors in the minimum pressure achieved are the thickness of the sacrificial wall relative to the characteristic dimension of the impacting particle, the density and material cohesion contrast of the sacrificial wall relative to the threat particle and the impact speed. The mass of the rear wall and the sacrificial wall are desirable to minimize for launch costs and dynamic concerns making it important to have an understanding of the effects of density contrast and impact speed. In this paper a fourth key parameter is identified related to fragmentation, which corresponds to the ratio of the size of the projectile relative to the transition from brittle to ductile hole growth in the projectile. Ballistic limit equations have been developed to define the failure limits of a MMOD shield, generally in terms of projectile diameter (or mass), impact velocity, and angle. Within the range of impact velocities relevant for Earth-orbiting spacecraft, three distinct regions of penetration phenomenology have been identified for Whipple shields: center dot Low velocity: the projectile is eroded (and possibly deformed) during its passage through the bumper plate, but is not fragmented. Thus, perforation of the rear wall is by a fragment with a mass and speed equal to or less than the original impactor. center dot Intermediate (shatter) velocity: impact velocities are sufficient to induce projectile fragmentation upon impact with the bumper plate, resulting in a coarse debris cloud with large solid fragments. Increasing velocity within the shatter regime results in increased fragmentation, and eventually melting, of the projectile and bumper fragments, generating a finer and more evenly dispersed debris cloud. Failure of the rear wall is a complicated combination of modes observed at low- and hypervelocity. center dot Hypervelocity: the projectile and holed-out bumper material is completely, or nearly completely, melted and/or vaporized by the initial impact. The resultant debris cloud impacts over a dispersed area of the rear wall, loading it impulsively and inducing failure through rupture or petalling. While each of these regimes are well observed with extensive empirical methods to describe these regions, differences in impactor materials, configurations of shields and questions about the limitations of the attainable impact speeds have left questions that are difficult to answer from completely empirical methods.

Miller, J. E.↗
NASA NTRS2011

Solar Dynamics Observatory Guidance, Navigation, and Control System Overview

The Solar Dynamics Observatory (SDO) was designed and built at the Goddard Space Flight Center, launched from Cape Canaveral on February 11, 2010, and reached its final geosynchronous science orbit on March 16, 2010. The purpose of SDO is to observe the Sun and continuously relay data to a dedicated ground station. SDO remains Sun-pointing throughout most of its mission for the instruments to take measurements of the Sun. The SDO attitude control system (ACS) is a single-fault tolerant design. Its fully redundant attitude sensor complement includes sixteen coarse Sun sensors (CSSs), a digital Sun sensor (DSS), three two-axis inertial reference units (IRUs), and two star trackers (STs). The ACS also makes use of the four guide telescopes included as a part of one of the science instruments. Attitude actuation is performed using four reaction wheels assemblies (RWAs) and eight thrusters, with a single main engine used to provide velocity-change thrust for orbit raising. The attitude control software has five nominal control modes, three wheel-based modes and two thruster-based modes. A wheel-based Safehold running in the attitude control electronics box improves the robustness of the system as a whole. All six modes are designed on the same basic proportional-integral-derivative attitude error structure, with more robust modes setting their integral gains to zero. This paper details the final overall design of the SDO guidance, navigation, and control (GN&C) system and how it was used in practice during SDO launch, commissioning, and nominal operations. This overview will include the ACS control modes, attitude determination and sensor calibration, the high gain antenna (HGA) calibration, and jitter mitigation operation. The Solar Dynamics Observatory mission is part of the NASA Living With a Star program, which seeks to understand the changing Sun and its effects on the Solar System, life, and society. To this end, the SDO spacecraft carries three Sun-observing instruments: Helioseismic and Magnetic Imager (HMI), led by Stanford University; Atmospheric Imaging Assembly (AIA), led by Lockheed Martin Space and Astrophysics Laboratory; and Extreme Ultraviolet Variability Experiment (EVE), led by the University of Colorado. The basic mission is to observe the Sun for a very high percentage of the 5-year mission (10-year goal) with long stretches of uninterrupted observations and with constant, high-data-rate transmission to a dedicated ground station to be located in White Sands, New Mexico. These goals guided the design of the spacecraft bus that will carry and service the three-instrument payload. Overarching design goals for the bus are geosynchronous orbit, near-constant Sun observations with the ability to fly through eclipses, and constant HGA contact with the dedicated ground station. A three-axis stabilized ACS is needed both to point at the Sun accurately and to keep the roll about the Sun vector correctly positioned with respect to the solar north pole. This roll control is especially important for the magnetic field imaging of HM I. The mission requirements have several general impacts on the ACS design. Both the AIA and HMI instruments are very sensitive to the blurring caused by jitter. Each has an image stabilization system (ISS) with some ability to filter out high frequency motion, but below the bandwidth of the ISS the control system must compensate for disturbances within the ACS bandwidth or avoid exciting jitter at higher frequencies. Within the ACS bandwidth, the control requirement imposed by AIA is to place the center of the solar disk no more than 2 arc sec, 3 , from a body-defined target based on one of the GTs that accompany the instrument. This body-defined target, called the science reference boresight (SRB), was determined from the postlaunch orientation of the GTs by averaging the bounding telescope boresights for pitch to get a pitch SRB coordinate, and by averaging the bounding boresights for yaw toet the yaw SRB coordinate. The location of this SRB in the 0.5-deg field-of-view for each GT then becomes the central target for each telescope; one GT is selected for use as the ACS controlling guide telescope (CGT) at any given time. Fine Sun-pointing is effected based on this SRB for all three instruments when the Sun is within the linear range of the CGT. In addition to limiting jitter, HMI science requires averaging several observations, making the instrument sensitive to low frequency motion that induces differential motion between each observation. This requires the spacecraft attitude to be stable about the roll axis to approximately 10 arcsec over a ten-minute period. Instrument calibrations require that the spacecraft point the SRB up to 2.5 degrees in pitch and yaw away from the center of the Sun, placing the Sun outside the field-of-view of the guide telescopes. In such instances, when the GTs cannot provide the definitive target for the ACS, on-board attitude determination combined with ephemeris prediction of the Sun direction must provide the definitive target. EVE is capable of observing the Sun with less dependence on attitude control. However, the ground data processing needs for calibrations result in the most strict attitude knowledge requirements for the mission: [35,70,70] arcsec, 3 , of knowledge with respect to the center of the solar disk. In addition to driving the ACS sensor selection, the knowledge requirements, which have their effect primarily during Inertial mode calibrations, drive the accuracy requirements for the solar ephemeris. The need to achieve and maintain geosynchronous orbit (GEO) drove the need for high-efficiency propulsive systems and appropriate attitude control. The main engine provided high specific impulse for the maneuvers to attain GEO, while the smaller ACS thrusters managed the disturbance torques of the larger engine and provided the capability for much smaller adjustment burns on orbit. SDO s large solar profile means that solar radiation pressure is a large torque disturbance, and the momentum buildup from this disturbance and the GEO altitude drives the ACS to use thrusters to manage vehicle momentum. The demanding data capture budget for the mission, however, requires SDO to avoid frequent thruster maneuvers, while concerns about on-orbit jitter restrict the maximum desired wheel speeds desired from the RWAs. The plan for on-orbit wheel speed and momentum management will be discussed as well as what is now being done in operation after the jitter environment was characterized. The SDO ACS hardware complement is single-fault tolerant. Two main processors carry virtually identical copies of the command and data handling and ACS software, and two identical attitude control electronics (ACE) boxes carry Coldfire processors with contingency ACS software and other hardware interface cards; the ACE structure allows reaction wheels to be commanded by the Sun-pointing Safehold independent of the Mil Std 1553 data bus. The sixteen Adcole CSSs are grouped into primary and backup sets of eight sensors, each set providing the ability to calculate a sun vector. Each set of eight eyes provides full 4 -steradian coverage. The Adcole DSS comprises an optics head and a separate electronics box providing a 1553 data interface. The electronics box is mounted inside the Faraday cage created by the spacecraft bus module. The DSS head with its 32- deg square FOV is mounted on the instrument module with its boresight along the spacecraft X axis, nearly aligned with the Sun during observations. Adcole has designed the DSS calibration parameters so that the accuracy is 0.24 arcminutes within 10 deg of the boresight, and diminishes to 3 arcminutes as the Sun moves towards the edges of its FOV . This DSS calibration scheme provides higher accuracy attitude determination over the range of the instrument calibration maneuvers.

Morgenstern, Wendy M.↗
NASA NTRS2018

HOMES: Holographic Method for Exoplanet Spectroscopy

The search for habitable exoplanets in the "neighborhood," that is within 10 parsecs (pc),invites new observational techniques, because telescopes available today have significant limitations. For example, gravitational micro-lensing is a deep field method that has little or no utility within 10 pc. Photometry of transits reduces the availability of exoplanets to a probability of approximately a 1%, and fewer than 1 out of 100 earth twins would ever be seen. For the 1% of exoplanets seen in transit, spectra of albedos cannot be taken.Even more limiting, Doppler shift studies are indirect and cannot characterize exoplanets other than by their orbits and masses through stellar radial velocity (RV). Moreover, RV by Doppler shift cannot detect planets in orbits on a plane perpendicular to our line of sight. Astrometry can be used to measure stellar orbital wobble for those exoplanetary systems in the plane perpendicular to our line of sight, as was contemplated for the cancelled NASA SIM mission 1 and rests now with ESA's GAIA2. In either case, the exoplanets detected by stellar RV or astrometry would not be directly observed, so the assumption that the three phases of water could exist does not answer a key question of habitability. Is there water on the exoplanet? Moreover, the discovery of water on an exoplanet in the habitable zone, exciting though it would be, is insufficient to determine just how habitable the exoplanet might possibly be. Like Goldilocks' porridge - there are many other ingredients beyond water alone in a life-sustaining soup.We propose a new species of optical telescope that has as one of its capabilities the high resolution spectrographic characterization of exoplanets on stars within 10 pc of the observatory. To achieve this performance specification, we propose to break with convention by taking advantage of an optical technology unknown until relatively recent times: the hologram. In this Report we show how holograms used in primary and secondary can extract spectrograms from exoplanets at distances reaching 33 light years from our solar system. We describe a notional space telescope using these novel optics that could make its observations in a space deployment scenario.

Ditto, Thomas↗
NASA NTRS2003

Solar Thermal Propulsion Improvements at Marshall Space Flight Center

Solar Thermal Propulsion (STP) is a concept which operates by transferring solar energy to a propellant, which thermally expands through a nozzle. The specific impulse performance is about twice that of chemical combustions engines, since there is no need for an oxidizer. In orbit, an inflatable concentrator mirror captures sunlight and focuses it inside an engine absorber cavity/heat exchanger, which then heats the propellant. The primary application of STP is with upperstages taking payloads from low earth orbit to geosynchronous earth orbit or earth escape velocities. STP engines are made of high temperature materials since heat exchanger operation requires temperatures greater than 2500K. Refractory metals such as tungsten and rhenium have been examined. The materials must also be compatible with hot hydrogen propellant. MSFC has three different engine designs, made of different refractory metal materials ready to test. Future engines will be made of high temperature carbide materials, which can withstand temperatures greater than 3000K, hot hydrogen, and provide higher performance. A specific impulse greater than 1000 seconds greatly reduces the amount of required propellant. A special 1 OkW solar ground test facility was made at MSFC to test various STP engine designs. The heliostat mirror, with dual-axis gear drive, tracks and reflects sunlight to the 18 ft. diameter concentrator mirror. The concentrator then focuses sunlight through a vacuum chamber window to a small focal point inside the STP engine. The facility closely simulates how the STP engine would function in orbit. The flux intensity at the focal point is equivalent to the intensity at a distance of 7 solar radii from the sun.

Gerrish, Harold P.↗
NASA NTRS1975

Close encounters of small bodies and planets

A small body in an orbit which crosses that of a planet may be eliminated by collision or ejected from the solar system. A method is developed for computing the probabilities of these fates. With some simplifying assumptions, the probabilities can be expressed as functions of the relative velocity only, rather than the orbital elements. The ejection probability does not vary greatly with velocity. The results are in agreement with those of Bandermann and Wolstencroft, but not with those of Opik; a reason for this disagreement is suggested. This method can be adapted to compute probabilities of generalized orbital changes. Some applications are discussed.

Weidenschilling, S. J.↗
NASA NTRS2023

A High-Eccentricity Warm Jupiter Orbiting TOI-4127

We report the discovery of TOI-4127 b, which is a transiting, Jupiter-sized exoplanet on a long period (P=56.39879 +0.00010 -0.00010 days) and a high-eccentricity orbit around a late F-type dwarf star. This warm Jupiter was first detected and identified as a promising candidate from a search for single-transit signals in TESS Sector 20 data and was later characterized as a planet following two subsequent transits (TESS Sectors 26 and 53) and follow-up ground-based RV observations with the NEID and SOPHIE spectrographs. We jointly fit the transit and RV data to constrain the physical (R p =1.096 +0.039 -0.032 R J , M p =2.30 +0.11 -0.11 M J ) and orbital parameters of the exoplanet. Given its high orbital eccentricity (𝓮=0.7471 +0.0078 -0.0086 ), TOI-4127 b is a compelling candidate for studies of warm Jupiter populations and of hot Jupiter formation pathways. We show that the present periastron separation of TOI-4127 b is too large for high-eccentricity tidal migration to circularize its orbit, and that TOI-4127 b is unlikely to be a hot Jupiter progenitor unless it is undergoing angular momentum exchange with an undetected outer companion. Although we find no evidence for an external companion, the available observational data are insufficient to rule out the presence of a perturber that can excite eccentricity oscillations and facilitate tidal migration.

Exoplanet astronomy↗
NASA NTRS1963

LOOK AT THE REENTRY PROBLEM

Reentry guidance and control in near-earth orbits and near escape velocity as well as in deep space and at velocities in excess of escape velocity

DEEP SPACE↗
NASA NTRS1971

The relative effects of entry parameters on thermal protection system weight

Shielding a spacecraft from the severe thermal environment of an atmospheric entry requires a sophisticated thermal protection system (TPS). Thermal computer program models were developed for two such TPS designs proposed for the space shuttle orbiter. The multilayer systems, a reusable surface insulation TPS, and a re-radiative metallic skin TPS, were sized for a cross-section of trajectories in the entry corridor. This analysis indicates the relative influence of the entry parameters on the weight of each TPS concept. The results are summarized graphically. The trajectory variables considered were down-range, cross-range, orbit inclination, entry interface velocity and flight path angle, maximum heating rate level, angle of attack, and ballistic coefficient. Variations in cross-range and flight path angle over the ranges considered had virtually no effect on the required entry TPS weight. The TPS weight was significantly more sensitive to variations in angle of attack than to dispersions in the other trajectory considered.

Hirasaki, P. N.↗
NASA NTRS1977

An altitude-dependent spacecraft charging model

A model for the altitude dependence of the hot plasma parameters responsible for the electrostatic charging of spacecraft was developed. Based upon plasma orbit theory, the directed velocity is a function of the ambient magnetic field flux density. A consequence of this approach is that while the thermal velocity distributions (assumed to be Maxwellian) of the plasma particles are independent of the magnetic field strength (and hence altitude), the particle densities increase with magnetic field strength. Thus, according to this model, while the equilibrium voltage is independent of altitude, the charging current density increases with decreasing altitude. However, the probability of such spacecraft charging decreases with decreasing altitude.

Haffner, J. W.↗