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The inability to test planetary spacecraft in the flight environment prior to a mission requires engineers to rely on ground-based testing and models of the vehicle and expected environments. One of the most widely used engineering models of the atmosphere is the Global Reference Atmospheric Model (GRAM) developed and maintained by the NASA Marshall Space Flight Center (MSFC). The NASA Science Mission Directorate (SMD) has provided funding support to upgrade the GRAMs.

The inability to test planetary spacecraft in the flight environment prior to a mission requires engineers to rely on ground-based testing and models of the vehicle and expected environments. One of the most widely used engineering models of the Martian atmosphere is the Mars Global Reference Atmospheric Model (Mars-GRAM) developed and maintained by the NASA Marshall Space Flight Center (MSFC). The NASA Science Mission Directorate (SMD) has provided funding support to upgrade the planetary GRAMs in Fiscal Year 2018 and 2019. This poster summarizes the upgrades that have been made to Mars-GRAM, the release status of Mars-GRAM, the new GRAMs that are under development, and future Mars-GRAM upgrade plans.

NASA is at the forefront of planetary exploration. The inability to test planetary space-craft in the flight environment prior to a mission requires engineers to rely on ground-based testing and models of the vehicle and expected environments. One of the most widely used engineering models of the atmosphere for many NASA projects is the Global Reference Atmospheric Model (GRAM) developed by the NASA Marshall Space Flight Center (MSFC). Over the past decade GRAM upgrades and maintenance have depended on inconsistent and waning project-specific support. Recently, the NASA Science Mission Directorate (SMD) has agreed to provide funding support in Fiscal Year 2018 and 2019 to upgrade the GRAMs.

The Exo-Brake is a simple, non-propulsive means of de-orbiting small payloads from orbital platforms such as the International Space Station (ISS). Two de-orbiting experiments with fixed surface area Exo-Brakes have been successfully conducted in the last two years on the TechEdSat-3 and -4 nano-satellite missions. The development of the free molecular flow aerodynamic data-base is presented in terms of angle of attack, projected front surface area variation, and altitude. Altitudes are considered ranging from the 400km ISS jettison altitude to 90km. Trajectory tools are then used to predict de-orbit/entry corridors with the inclusion of the key atmospheric and geomagnetic uncertainties. Control system strategies are discussed which will be applied to the next two planned TechEdSat-5 and -6 nano-satellite missions - thus increasing the targeting accuracy at the Von Karman altitude through the proposed drag modulation technique.

A NASA flagship mission to Titan, the largest moon of Saturn and the only moon in the solar system with a significant atmosphere, has been designed that uses three separate spacecraft, each requiring significant interaction with the atmosphere. The first vehicle is a Titan lander for lower-atmosphere and surface science. The second is an aerial vehicle for aerial science at approximately 10 km altitude with an expected lifetime of one year. This spacecraft will use the natural winds of Titan to cover a large area over its lifetime. The third vehicle is a Titan orbiter that will interact with the atmosphere in two respects. The first atmospheric interaction is the orbital insertion maneuver that will be accomplished using aerocapture, during which time the hyperbolic approach of 6.5 km/s will be reduced to 1.6 km/s over 41 minutes with an exit periapsis altitude of 130 km. The second atmospheric interaction occurs after a propulsive maneuver has raised the periapsis after aerocapture to 1170 km, where the atmosphere will be sampled over several months. This is the first phase of aerosampling that covers southern latitudes. After a 3.3-year circular science phase at an altitude of 1700 km, a second phase of additional aerosampling is performed sampling northern latitudes. The atmospheric trajectory analysis for these three spacecraft will be discussed throughout this paper.

Within the present budgetary environment, developing the technology that leads to an operationally efficient space transportation system with the required performance is a challenge. The present research focuses on a methodology to determine high payoff technology investment strategies. Research has been conducted at Langley Research Center in which design codes for the conceptual analysis of space transportation systems have been integrated in a multidisciplinary design optimization approach. The current study integrates trajectory, propulsion, weights and sizing, and cost disciplines where the effect of technology maturation on the development cost of a single stage to orbit reusable launch vehicle is examined. Results show that the technology investment prior to full-scale development has a significant economic payoff. The design optimization process is used to determine strategic allocations of limited technology funding to maximize the economic payoff.

The investigation focuses on development of a rapid multidisciplinary analysis and optimization capability for launch-vehicle design. Two multidisciplinary optimization strategies in which the analyses are integrated in different manners are implemented and evaluated for solution of a single-stage-to-orbit launch-vehicle design problem. Weights and sizing, propulsion, and trajectory issues are directly addressed in each optimization process. Additionally, the need to maintain a consistent vehicle model across the disciplines is discussed. Both solution strategies were shown to obtain similar solutions from two different starting points. These solutions suggests that a dual-fuel, single-stage-to-orbit vehicle with a dry weight of approximately 1.927 x 10(exp 5)lb, gross liftoff weight of 2.165 x 10(exp 6)lb, and length of 181 ft is attainable. A comparison of the two approaches demonstrates that treatment or disciplinary coupling has a direct effect on optimization convergence and the required computational effort. In comparison with the first solution strategy, which is of the general form typically used within the launch vehicle design community at present, the second optimization approach is shown to he 3-4 times more computationally efficient.

Mars trajectory design options were examined that would accommodate a premature termination of a nominal manned opposition class mission for opportunities between 2010 and 2025. A successful abort must provide a safe return to Earth in the shortest possible time consistent with mission constraints. In this study, aborts that provided a minimum increase in the initial vehicle mass in low Earth orbit (IMLEO) were identified by locating direct transfer nominal missions and nominal missions including an outbound or inbound Venus swing-by that minimized IMLEO. The ease with which these missions could be aborted while meeting propulsion and time constraints was investigated by examining free return (unpowered) and powered aborts. Further reductions in trip time were made to some aborts by the addition or removal of an inbound Venus swing-by. The results show that, although few free return aborts met the specified constraints, 85% of each nominal mission could be aborted as a powered abort without an increase in propellant. Also, in many cases, the addition or removal of a Venus swing-by increased the number of abort opportunities or decreased the total trip time during an abort.

A study has been conducted of the staging maneuver for a two-stage, vertical-takeoff, launch vehicle and the subsequent glideback of the booster to a launch site runway. The booster and orbiter are both winged, fully reusable, and have liquid rocket main propulsion. for the staging maneuver, which nominally occurs at Mach 3, a flight control system and separation technique were devised which produced successful separations over a wide range of staging angles of attack. A guidance algorithm was developed for the unpowered glide-back of the booster from the completion of the staging maneuver to touchdown on a runway in the vicinity of the launch site. Using this guidance technique, the booster successfully touched down on the runway while being subjected to a wide range of seasonal and perturbed atmospheric conditions with and without wids as well as errors in staging state conditions and errors in predicted aerodynamics.

The ascent, abort, and entry capability of a vehicle for rescue of the Space Station crew or ferry of personnel and supplies to and from the Space Station are analyzed. The configurations of the Space Station rescue vehicle and the Space Station personnel/logistics vehicle (SSPLV) are discussed and illustrated. The nominal ascent trajectory for the SSPLV delivered to orbit on a Titan II is presented. The ascent abort modes from launch to orbital injection are evaluated. It is shown that five landing sites with runways longer than 10,000 ft could provide a landing opportunity from every orbit of the Space Station in a 220-nmi circular orbit with a 28.5 deg inclination.

The preliminary design of a horizontal takeoff, horizontal landing, single-stage-to-orbit, rocket-powered space vehicle was performed. The purpose was to examine technology requirements for future small payload launch vehicles. The distinguishing aspect of the design was the utilization of thrust vectoring to provide half of the lift at takeoff. The inclusion of a canard was necessary to provide additional lift at takeoff and to balance the moments produced with thrust vectoring. A weights estimation, an aerodynamic assessment, a trajectory analysis, and a gear weight analysis were performed. The takeoff weight of the resulting vehicle was approximately 1.26 million pounds, based on advanced technology structures and subsystems. The vehicle was designed to deliver a 5000-pound payload to a polar orbit.

The aerodynamic database upon which the missions and the proposed designs of an atmospheric entry research vehicle (ERV) would be based are discussed, along with a trajectory analysis of two missions proposed for the ERV. The ERV is intended as a means to explore reentry techniques, thermal protection systems and cross-range and orbital change maneuvers beyond the operational envelope of the Orbiter. The 25 ft long ERV would have a 13 ft wingspan, allowing it to be launched from the Orbiter bay. The ERV would have atmospheric maneuvering control surfaces and reaction control systems for orbital maneuvers. The total weight would be 12,000 lb. The aerodynamic design for the ERV was developed from wind tunnel, theoretical and Shuttle data for continuum, transition and free-molecule flight regimes. Typical flight profiles for intended missions are outlined.

A slant-nose-cylinder aeroassisted orbital transfer vehicle configuration is described and analyzed in this study. The vehicle is sized for a 12,000 lb roundtrip payload between low earth orbit and geosynchronous orbit and is assumed to be space based. The vehicle can be fabricated using near-term technologies and is fully reusable. Optional advanced technologies offer potential for improved performance. The vehicle can be assembled on the ground and carried to orbit in the Shuttle cargo bay. An enclosed payload bay is provided in the vehicle to protect payloads during the pass through the atmosphere. The payload bay capacity can be increased from a 10 ft to a 14 ft diameter payload by replacing a modular section of the payload bay in space. The results of calculations used to size the vehicle and to predict its performance and weight are presented.

Within the last year, there has been a growing interest in technology which would support the development of vehicles which can maneuver in the atmosphere while returning from orbit and vehicles which can fly in the atmosphere at hypersonic speeds for sustained periods of time. This rebirth of interest in hypersonics is partly a result of a developing awareness of the potential benefits which can be derived from vehicles with capability to operate between the limits of existing aircraft and spacecraft. Examples of the types of maneuvers which are projected for this new class of vehicles are presented. Challenges provided by the considered maneuvers are summarized, taking into account a synergetic or atmospheric plane change. The plane change maneuver will require a Thermal Protection System (TPS). Attention is given to the definition of a Shuttle launched entry research vehicle experiment. Details regarding the synergetic plane change are considered along with the maneuvering entry, vehicle heating, and experiments and instrumentation.

Factors affecting the bailout characteristics from the space shuttle orbiter at low-subsonic speeds were investigated. In the 12-foot low-speed tunnel and the 4 by 7-meter tunnel with 0.03-scale models. The effect of crew-model exit velocity, body position, and body weight were studied with egress from the main side hatch with the orbiter upright and from the upper cabin hatch with the orbiter inverted. Crew model drag and flow field measurements around the orbiter were estimated. The high-angle-of-attack trim characteristics of the orbiter was determined by force tests in an attempt to improve bailout conditions. A computer simulation was made to evaluate the maneuver necessary to attain the high-angle-of-attack trim.

A six degree of freedom simulation analysis was performed for the space shuttle orbiter during entry from Mach 8 to Mach 1.5 with realistic off nominal conditions by using the flight control systems defined by the shuttle contractor. The off nominal conditions included aerodynamic uncertainties in extrapolating from wind tunnel derived characteristics to full scale flight characteristics, uncertainties in the estimates of the reaction control system interaction with the orbiter aerodynamics, an error in deriving the angle of attack from onboard instrumentation, the failure of two of the four reaction control system thrusters on each side, and a lateral center of gravity offset coupled with vehicle and flow asymmetries. With combinations of these off nominal conditions, the flight control system performed satisfactorily. At low hypersonic speeds, a few cases exhibited unacceptable performances when errors in deriving the angle of attack from the onboard instrumentation were modeled. The orbiter was unable to maintain lateral trim for some cases between Mach 5 and Mach 2 and exhibited limit cycle tendencies or residual roll oscillations between Mach 3 and Mach 1. Piloting techniques and changes in some gains and switching times in the flight control system are suggested to help alleviate these problems.

Aero-assisted orbital transfer vehicles are analyzed. The aerodynamic characteristics over the flight profile and three- and six-degree-of-freedom performance analyses were determined. The important results, to date, are: (1) the aerodynamic preliminary analysis system, an interactive computer program, used to predict the aerodynamics (performance, stability, and control) for these vehicles; (2) the performance capability, e.g., maximum inclination change, maximum heating rate, and maximum sensed acceleration, can be determined using continuum aerodynamics only; (3) guidance schemes can be developed that allow for errors in atmospheric density prediction, mispredicted trim angle of attack, and off-nominal atmospheric interface conditions, even for vehicles with a low lift-to-drag ratio; and (4) multiple pass trajectories can be used to reduce the maximum heating rate.

A six-degree-of-freedom simulation analysis has been performed for the Space Shuttle Orbiter during entry from Mach 10 to 2.5 with realistic off-nominal conditions using the entry flight control system specified in May 1978. The off-nominal conditions included the following: (1) aerodynamic uncertainties, (2) an error in deriving the angle of attack from onboard instrumentation, (3) the failure of two of the four reaction control-system thrusters on each side, and (4) a lateral center-of-gravity offset. With combinations of the above off-nominal conditions, the control system performed satisfactorily with a few exceptions. The cases that did not exhibit satisfactory performance displayed the following main weaknesses. Marginal performance was exhibited at hypersonic speeds with a sensed angle-of-attack error of 4 deg. At supersonic speeds the system tended to be oscillatory, and the system diverged for several cases because of the inability to hold lateral trim. Several system modifications were suggested to help solve these problems and to maximize safety on the first flight: alter the elevon-trim and speed-brake schedules, delay switching to rudder trim until the rudder effectiveness is adequate, and reduce the overall rudder loop gain. These and other modifications were incorporated in a flight-control-system redesign in May 1979.

A feasibility study to assess the benefits/penalties of installing tip-fin controllers on the Space Shuttle orbiter has been conducted. The results from the aerodynamic and stability and control assessment of this study are presented. These results show that using tip-fin controllers with the centerline vertical tail removed provides adequate control for entry. The orbiter basic maneuvering capability is maintained, and symmetrical deflection of the controllers provides adequate speedbrake authority for energy management. Gust analyses of the orbiter with the tip-fin controller configuration have demonstrated no compromise of existing cross-wind landing capability. Six-degree-of-freedom simulations have shown the tip-fin controllers can provide adequate control for entry with no degradation in vehicle capability.

Studies are underway to design reusable orbital transfer vehicles that would be used to transfer payloads from low-earth orbit to higher orbits and return. One promising concept is to use an atmospheric pass on the return leg to reduce the amount of fuel for the mission. This paper discusses a six-degree-of-freedom simulation analysis for two configurations, a low-lift-to-drag ratio configuration and a medium-lift-to-drag ratio configuration using both a predictive guidance technique and an adaptive guidance technique. Both guidance schemes were evaluated using the 1962 standard atmosphere and three atmospheres that had been derived from three entries of the Space Shuttle. The predictive technique requires less reaction control system activity for both configurations, but because of the limited number of updates and because each update used the 1962 standard atmosphere, the adaptive technique produces more accurate exit conditions.

The Future Space Transportation System (FSTS) study was conducted by the NASA Langley Research Center to identify the technology requirements for concepts that will replace the Space Shuttle in the post 2000 time frame. The configuration chosen for the study is a two-stage, fully reusable, vertical liftoff, and horizontal landing system with a 150,000 lb. payload capability. The two stages are burned in parallel with the booster providing all the propellant until staging, which results in a large lateral c.g. movement. Nominally, the booster stages at Mach 3 and glides back to the launch site. Because of the large lateral c.g. travel, a scheme to trim the vehicle until staging occurred was developed that used both gimballing and throttling of the engines. Preliminary booster aerodynamics were determined, and the booster glideback trajectory was analyzed with and without winds. Finally, a preliminary abort analysis was conducted for each stage.

In the recent Future Space Transportation System (FSFS) study, a mission model was selected and a baseline vehicle which best met model requirements was shaped. This baseline was then analyzed for flight performance, structural and subsystem weight, and operation. Figures related to a payload of 150,000 lb in a 20-ft-diam by 90-ft-long envelope became the baseline. The existence of both space-based orbital transfer vehicles (OTVs) and a space station was assumed, taking into account a transfer of the payload from the launch vehicle to OTVs at the space station for final delivery to geosynchronous orbit (GEO). A computer-aided engineering system called Aerospace Vehicle Interactive Design (AVID) was employed in connection with baseline vehicle development. It was found that approximately three-fifth of the payload weight would be cryogenic propellants for OTVs. Attention is given to problems regarding the packaging of cryogenic tankage, a payload shroud, and studies of staging for two different booster propulsion units.

Flight data obtained from three Space Transportation System orbiter entries (STS-1, 2, and 3) are processed and analyzed to determine the roll interactions caused by the firing of the entry reaction control system (RCS). Comparisons between the flight-derived parameters and the predicted derivatives without interaction effects are made. The flight-derived RCS Plume flow-field interaction effects are independently deduced by direct integration of the incremental changes in the wing upper surface pressures induced by RCS side thruster firings. The separately obtained interaction effects are compared to the predicted values and the differences are discussed.

Aerodynamic and flight control analyses have shown that the application of the tip-fin controller and removal of the centerline vertical tail does not produce improved flyability of the Shuttle Orbiter in the supersonic speed regime. Preliminary design studies show that removal of the centerline vertical tail and the installation of tip-fin controllers could result in savings up to 900 kg. It is also shown that the reaction control system could be deactivated much sooner than it is possible with the present nominal orbiter configuration.

A nonlinear, six degree of freedom, digital computer simulation of a vehicle which has constant mass properties and whose attitudes are controlled by both aerodynamic surfaces and reaction control system thrusters was developed. A rotating, oblate Earth model was used to describe the gravitational forces which affect long duration Earth entry trajectories. The program is executed in a nonreal time mode or connected to a simulation cockpit to conduct piloted and autopilot studies. The program guidance and control software used by the space shuttle orbiter for its descent from approximately 121.9 km to touchdown on the runway.

An approach has been developed to determine the guidance and control system sensitivity to off-nominal aerodynamics for the Space Shuttle Orbiter during entry. This approach, which uses a nonlinear six-degree-of-freedom interactive, digital simulation, has been applied to both the longitudinal and lateral-directional axes for a portion of the orbiter entry. Boundary values for each of the aerodynamic parameters have been identified, the key parameters have been determined, and system modifications that will increase system tolerance to off-nominal aerodynamics have been recommended. The simulations were judged by specified criteria and the performance was evaluated by use of key dependent variables. The analysis is now being expanded to include the latest shuttle guidance and control systems throughout the entry speed range.

There are six degree-of-freedom simulations of the space shuttle orbiter entry with aerodynamic control hysteresis conducted on the NASA Langley Research Center interactive simulator known as the Automatic Reentry Flight Dynamics Simulator. These were performed to determine if the presence of aerodynamic control hysteresis would endanger the mission, either by making the vehicle unable to maintain proper attitude for a safe entry, or by increasing the amount of the reaction control system's fuel consumption beyond that carried.

An investigation has been made of abort procedures for space shuttle-type vehicles using a point mass trajectory optimization program known as POST. This study determined the minimum time gap between immediate and once-around safe return to the launch site from a baseline due-East launch trajectory for an alternate space shuttle concept which experiences an instantaneous loss of 25 percent of the total main engine thrust.

A six-degree-of-freedom simulation analysis was performed for the space shuttle orbiter entry from Mach 10 to Mach 2.5 with realistic off-nominal conditions using the flight control system referred to as the November 1976 Integrated Digital Autopilot. The off-nominal conditions included: (1) aerodynamic uncertainties in extrapolating from wind tunnel of flight characteristics, (2) error in deriving angle of attack from onboard instrumentation, (3) failure of two of the four reaction control-system thrusters on each side (design specification), and (4) lateral center-of-gravity offset. Many combinations of these off-nominal conditions resulted in a loss of the orbiter. Control-system modifications were identified to prevent this possibility.