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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS1968

Description and Flight Performance Results of the WASP Sounding Rocket

A general description of the design and construction of the WASP sounding rocket and of the performance of its first flight are presented. The purpose of the flight test was to place the 862-pound (391-kg) spacecraft above 250 000 feet (76.25 km) on free-fall trajectory for at least 6 minutes in order to study the effect of "weightlessness" on a slosh dynamics experiment. The WASP sounding rocket fulfilled its intended mission requirements. The sounding rocket approximately followed a nominal trajectory. The payload was in free fall above 250 000 feet (76.25 km) for 6.5 minutes and reached an apogee altitude of 134 nautical miles (248 km). Flight data including velocity, altitude, acceleration, roll rate, and angle of attack are discussed and compared to nominal performance calculations. The effect of residual burning of the second stage motor is analyzed. The flight vibration environment is presented and analyzed, including root mean square (RMS) and power spectral density analysis.

De Pauw, J. F.↗
NASA NTRS1992

Lunar scout: A Project Artemis proposal

The results of a student project to design a lunar lander in the context of a specifically defined mission are presented. The Lunar Scout will be launched from Cape Canaveral, Florida onboard a Delta II launch vehicle. The Delta II will carry the lander and its payload to a 1367 km orbit. Once it reaches that altitude, a STAR 48A solid rocket motor will kick the spacecraft into a lunar trajectory. After burnout of the lunar insertion motor, it will be jettisoned from the spacecraft. The flight from the earth to the moon will take approximately 106.4 hours. During this time the battery, which was fully charged prior to launch, will provide all power to the spacecraft. Every hour, the spacecraft will use its sun sensors and star trackers to update its position, maintain some stabilization and relay it back to earth using the dipole antennas. At the start of its lunar trajectory, the spacecraft will fire one of its 1.5 N thrusters to spin in at a very small rate. The main reason for this is to prevent one side of the spacecraft from overheating in the sun. When the spacecraft nears the moon, it will orient itself for the main retro burn. At an altitude of 200 km, a 4400 N bipropellant liquid thruster will ignite to slow the spacecraft. During the burn, the radar altimeter will be turned on to guide the spacecraft. The main retro rocket will slow the lander to 10 m/s at an approximate altitude of 40 km above the moon. From there, the space craft will use four 4.5 N hydrazine vertical thrusters and 1.5 N horizontal thrusters to guide the spacecraft to a soft landing. Once on the ground, the lander will shutoff the radar and attitude control systems. After the debris from the impact has settled, the six solar panels will be deployed to begin recharging the batteries and to power up the payload. The feedhorn antenna will then rotate to fix itself on the earth. Once it moves, it will stay in that position for the spacecraft's lifetime. The payload will then be activated to begin the lunar mission.

Source record↗
NASA NTRS1972

Rapid optimization of multiple-burn rocket flights.

Different formulations of the fuel optimization problem for multiple burn trajectories are considered. It is shown that certain customary idealizing assumptions lead to an ill-posed optimization problem for which no solution exists. Several ways are discussed for avoiding such difficulties by more realistic problem statements. An iterative solution of the boundary value problem is presented together with efficient coast arc computations, the right end conditions for various orbital missions, and some test results.

Brown, K. R.↗
NASA NTRS2011

The Application of Lean Thinking Principles and Kaizen Practices for the Successful Development and Implementation of the Ares I-X Flight Test Rocket and Mission

On October 28, 2009 the Ares I-X flight test rocket launched from Kennedy Space Center and flew its suborbital trajectory as designed. The mission was successfully completed as data from the test, and associated development activities were analyzed, transferred to stakeholders, and well documented. Positive lessons learned from Ares I-X were that the application of lean thinking principles and kaizen practices are effective in streamlining development activities. Ares I-X, like other historical rocket development projects, was hampered by technical, cost, and schedule challenges and if not addressed boldly could have resulted in cancellation of the test. The mission management team conducted nine major meetings, referred to as lean events, across its elements to assess plans, procedures, processes, requirements, controls, culture, organization, use of resources, and anything that could be changed to optimize schedule or reduce risk. The preeminent aspect of the lean events was the focus on value added activities and the removal or at least reduction in non-value activities. Trained Lean Six Sigma facilitators assisted the Ares I-X developers in conducting the lean events. They indirectly helped formulate the mission s own unique methodology for assessing schedule. A core team was selected to lead the events and report to the mission manager. Each activity leveraged specialized participants to analyze the subject matter and its related processes and then recommended alternatives and solutions. Stakeholders were the event champions. They empowered and encouraged the team to succeed. The keys to success were thorough preparation, honest dialog, small groups, adherence to the Ares I-X ground rules, and accountability through disciplined reporting and tracking of actions. This lean event formula was game-changing as demonstrated by the success of Ares I-X. It is highly recommended as a management tool to help develop other complex systems efficiently. The key benefits for Ares I-X were obtaining unambiguous schedule margin, defining enabling options for risk reduction, and most importantly a stronger more unified team.

Askins, B. R.↗
NASA NTRS2008

Lagrangian Trajectory Modeling of Lunar Dust Particles

Apollo landing videos shot from inside the right LEM window, provide a quantitative measure of the characteristics and dynamics of the ejecta spray of lunar regolith particles beneath the Lander during the final 10 [m] or so of descent. Photogrammetry analysis gives an estimate of the thickness of the dust layer and angle of trajectory. In addition, Apollo landing video analysis divulges valuable information on the regolith ejecta interactions with lunar surface topography. For example, dense dust streaks are seen to originate at the outer rims of craters within a critical radius of the Lander during descent. The primary intent of this work was to develop a mathematical model and software implementation for the trajectory simulation of lunar dust particles acted on by gas jets originating from the nozzle of a lunar Lander, where the particle sizes typically range from 10 micron to 500 micron. The high temperature, supersonic jet of gas that is exhausted from a rocket engine can propel dust, soil, gravel, as well as small rocks to high velocities. The lunar vacuum allows ejected particles to travel great distances unimpeded, and in the case of smaller particles, escape velocities may be reached. The particle size distributions and kinetic energies of ejected particles can lead to damage to the landing spacecraft or to other hardware that has previously been deployed in the vicinity. Thus the primary motivation behind this work is to seek a better understanding for the purpose of modeling and predicting the behavior of regolith dust particle trajectories during powered rocket descent and ascent.

Lane, John E.↗
NASA NTRS2011

The Application of Lean Thinking Principles and Kaizen Practices for the Successful Development and Implementation of the Ares I-X Flight Test Rocket and Mission

On October 28, 2009 the Ares I-X flight test rocket launched from Kennedy Space Center and flew its suborbital trajectory as designed. The mission was successfully completed as data from the test, and associated development activities were analyzed, transferred to stakeholders, and well documented. A positive lesson learned from Ares I-X was that the application of lean thinking principles and kaizen practices was very effective in streamlining development activities. Ares I-X, like other historical rocket development projects, was hampered by technical, cost, and schedule challenges and if not addressed boldly could have resulted in cancellation of the test. The mission management team conducted nine major meetings, referred to as lean events, across its elements to assess plans, procedures, processes, requirements, controls, culture, organization, use of resources, and anything that could be changed to optimize schedule or reduce risk. The preeminent aspect of the lean events was the focus on value added activities and the removal or at least reduction in non-value added activities. Trained Lean Six Sigma facilitators assisted the Ares I-X developers in conducting the lean events. They indirectly helped formulate the mission s own unique methodology for assessing schedule. A core team was selected to lead the events and report to the mission manager. Each activity leveraged specialized participants to analyze the subject matter and its related processes and then recommended alternatives and solutions. Stakeholders were the event champions. They empowered and encouraged the team to succeed. The keys to success were thorough preparation, honest dialog, small groups, adherence to the Ares I-X ground rules, and accountability through disciplined reporting and tracking of actions. This lean event formula was game-changing as demonstrated by Ares I-X. It is highly recommended as a management tool to help develop other complex systems efficiently. The key benefits for Ares I-X were obtaining unambiguous schedule margin, defining enabling options for risk reduction, and most importantly a stronger more unified team.

Askins, B. R.↗
NASA NTRS2011

Overview of Experimental Investigations for Ares I Launch Vehicle Development

Another concern for the vehicle during its design trajectory was the separation of the first stage solid rocket booster from the upper stage component after it had depleted its solid fuel propellant. There has been some concern about the interstage of the first stage from clearing the nozzle of the J2-X engine. A detailed separation aerodynamic wind tunnel investigation was conducted in the AEDC VKF Tunnel A to help to investigate the interaction aerodynamic effects5. A comparison of the separation plane details between the Ares I architecture and the Ares I-X demonstration flight architecture is shown in figure 12. The Ares I design requires a more complex separation sequence and requires better control in order to avoid contact with the nozzle of the upper stage engine. The interstage, which houses the J2-X engine for the Ares I vehicle, must be able to separate cleanly to avoid contact of the J2-X engine. There is only about approximately 18 inches of buffer inside the interstage on each size of the nozzle so this is a challenging controlled separation event. This complex experimental investigation required two separate Ares I models (upper stage and first stage with interstage attached) with independent strain gauge balances installed in each model. It also required the Captive Trajectory System (CTS) that was needed to precisely locate the components in space relative to each other to fill out the planned test matrix. The model setup in the AEDC VKF Tunnel A is shown in figure 13. The CTS remotely positioned the first stage at the required x, y, and z positions and was able to provide interactions within 0.2" of the upper stage. A sample of the axial force on the first stage booster is shown in figure 14. These results, as a function of separation distance between the two stages, are compared to pre-test CFD results. Since this is a very challenging, highly unsteady flow field for CFD to correctly model, the experimental results have been utilized by GN&C discipline to more accurately represent the interaction aerodynamics. In addition to the integrated forces and moments obtained from the test, flow visualization data was obtained from this test in the form of Schlieren photographs, as shown in figure 15, which show the shock structure and interaction effects after the two stages separate during flight. This separation test was crucial in the successful flight test of the Ares I-X vehicle and provided the GN&C discipline with the unpowered proximity aerodynamic effect for a separation of the Ares I vehicle.

Tomek, William G.↗
NASA NTRS2000

Chance Encounter with a Stratospheric Kerosene Rocket Plume From Russia Over California

A high-altitude aircraft flight on April 18, 1997 detected an enormous aerosol cloud at 20 km altitude near California (37 N). Not visually observed, the cloud had high concentrations of soot and sulfate aerosol, and was over 180 km in horizontal extent. The cloud was probably a large hydrocarbon fueled vehicle, most likely from rocket motors burning liquid oxygen and kerosene. One of two Russian Soyuz rockets could have produced the cloud: a launch from the Baikonur Cosmodrome, Kazakhstan on April 6; or from Plesetsk, Russia on April 9. Parcel trajectories and long-lived trace gas concentrations suggest the Baikonur launch as the cloud source. Cloud trajectories do not trace the Soyuz plume from Asia to North America, illustrating the uncertainties of point-to-point trajectories. This cloud encounter is the only stratospheric measurement of a hydrocarbon fuel powered rocket.

Newman, P. A.↗
NASA NTRS1998

Rocketdyne Advanced Reusable Technologies Program Update

NASA Marshall Space Flight Center Advanced Reusable Technologies/Rocket Based Combined Cycle (RBCC) program has enabled: (1) the creation of unique trajectory simulation facility; (2) development of basic operability and performance data; (3) results to date: (a) rocket thrusters performing as designed (P(sub c), MR, q, etc); (b) AAR performance data measured at M = 0 and 3; (c) rocket mode performance at simulated vacuum conditions; and (d) data reduction in progress; (4) completing testing; M = 0, 3 and trajectory M = 3 to 3.7; and (5) next logical steps: build flight type engine and flight demonstrate RBCC.

Ratekin, Gary↗
NASA NTRS1993

Advanced launch system trajectory optimization using suboptimal control

The maximum-final mass trajectory of a proposed configuration of the Advanced Launch System is presented. A model for the two-stage rocket is given; the optimal control problem is formulated as a parameter optimization problem; and the optimal trajectory is computed using a nonlinear programming code called VF02AD. Numerical results are presented for the controls (angle of attack and velocity roll angle) and the states. After the initial rotation, the angle of attack goes to a positive value to keep the trajectory as high as possible, returns to near zero to pass through the transonic regime and satisfy the dynamic pressure constraint, returns to a positive value to keep the trajectory high and to take advantage of minimum drag at positive angle of attack due to aerodynamic shading of the booster, and then rolls off to negative values to satisfy the constraints. Because the engines cannot be throttled, the maximum dynamic pressure occurs at a single point; there is no maximum dynamic pressure subarc. To test approximations for obtaining analytical solutions for guidance, two additional optimal trajectories are computed: one using untrimmed aerodynamics and one using no atmospheric effects except for the dynamic pressure constraint. It is concluded that untrimmed aerodynamics has a negligible effect on the optimal trajectory and that approximate optimal controls should be able to be obtained by treating atmospheric effects as perturbations.

Shaver, Douglas A.↗
NASA NTRS1992

Orbital motion under continuous tangential thrust

The effect of continuous tangential thrust on the orbital motion and mass loss of a vehicle initially in a circular orbit is investigated analytically. It is shown that, for a thrust-to-weight ratio of greater than 0.16175, escape speed will eventually be reached along an unwinding spiral trajectory. For lower thrust-to-weight ratios, escape speed is never attained, and the flight path oscillates around a logarithmic spiral trajectory. Formulas are obtained for the approximate orbital motion and time of flight along each type of trajectory and for mass loss due to expenditure of rocket propellant.

Boltz, Frederick W.↗
NASA NTRS2007

Launching Payloads Into Orbit at Relatively Low Cost

A report proposes the development of a system for launching payloads into orbit at about one-fifth the cost per unit payload weight of current systems. The PILOT system was a solid-fuel, aerodynamically spun and spin-stabilized, five-stage rocket with onboard controls including little more than an optoelectronic horizon sensor and a timer for triggering the second and fifth stages, respectively. The proposal calls for four improvements over the PILOT system to enable control of orbital parameters: (1) the aerodynamic tipover of the rocket at the top of the atmosphere could be modeled as a nonuniform gyroscopic precession and could be controlled by selection of the initial rocket configuration and launch conditions; (2) the attitude of the rocket at the top of the first-stage trajectory could be measured by use of radar tracking or differential Global Positioning System receivers to determine when to trigger the second stage; (3) the final-stage engines could be configured around the payload to enhance spin stabilization during a half-orbit coast up to apoapsis where the final stage would be triggered; and (4) the final payload stage could be equipped with a "beltline" of small thrusters for correcting small errors in the trajectory as measured by an off-board tracking subsystem.

Wilcox, Brian↗
NASA NTRS2010

COMPASS Final Report: Low Cost Robotic Lunar Lander

The COllaborative Modeling for the Parametric Assessment of Space Systems (COMPASS) team designed a robotic lunar Lander to deliver an unspecified payload (greater than zero) to the lunar surface for the lowest cost in this 2006 design study. The purpose of the low cost lunar lander design was to investigate how much payload can an inexpensive chemical or Electric Propulsion (EP) system deliver to the Moon s surface. The spacecraft designed as the baseline out of this study was a solar powered robotic lander, launched on a Minotaur V launch vehicle on a direct injection trajectory to the lunar surface. A Star 27 solid rocket motor does lunar capture and performs 88 percent of the descent burn. The Robotic Lunar Lander soft-lands using a hydrazine propulsion system to perform the last 10% of the landing maneuver, leaving the descent at a near zero, but not exactly zero, terminal velocity. This low-cost robotic lander delivers 10 kg of science payload instruments to the lunar surface.

McGuire, Melissa L.↗
NASA NTRS2014

Passive Thermal Control for the Low Density Supersonic Decelerator (LDSD) Test Vehicle Spin Motors Sub-System

Future missions to Mars will require improved entry, descent, and landing (EDL) technology over the Viking-heritage systems which recently landed the largest payload to date, the 900 kg Mars Science Laboratory. As a result, NASA's Low Density Supersonic Decelerator (LDSD) project is working to advance the state of the art in Mars EDL systems by developing and testing three key technologies which will enable heavier payloads and higher altitude landing sites on the red planet. These technologies consist of a large 33.5 m diameter Supersonic Disk Sail (SSDS) parachute and two different Supersonic Inflatable Aerodynamic Decelerator (SIAD) devices - a robotic class that inflates to a 6 m diameter torus (SIAD-R), and an exploration class that inflates to an 8 m diameter isotensoid (SIADE). All three technologies will be demonstrated on test vehicles at high earth altitudes in order to simulate the Mars EDL environment. Each vehicle will be carried to altitude by a large helium balloon, released, spun up using spin motors to stabilize the vehicle's trajectory, and accelerated to supersonic speeds using a large solid rocket motor. The vehicle will then be spun down using another set of spin motors, and will deploy either the SIAD-R or SIAD-E, followed by the SSDS parachute until the vehicle lands in the ocean. Component level testing and bounding analysis are used to ensure the survival of system components in extreme thermal environments and predict temperatures throughout the flight. This paper presents a general description of the thermal testing, model correlation, and analysis of the spin motor passive thermal control sub-system to maintain spin motor performance, prescribed vehicle trajectory, and structural integrity of the test vehicle. The spin motor subsystem is predicted to meet its requirements with margin.

thermal analysis↗
NASA NTRS1977

Three-dimensional random earth atmospheres for Monte Carlo trajectory analyses

A set of four computer tapes containing random three dimensional Earth atmospheres is available for Monte Carlo trajectory analyses. The tapes contain sufficient atmospheric tables to allow replications of any trajectory below an altitude of 99 km. The atmospheres were provided by an empirical model designed to generate random atmospheres whose distributions match those in a data base of sounding rocket measurements. A readily implementable means of linking the tapes to any existing trajectory simulation computer program is described involving the addition of three subroutines which are listed in an appendix.

Campbell, J. W.↗
NASA NTRS1998

The Strutjet Rocket Based Combined Cycle Engine

The multi stage chemical rocket has been established over many years as the propulsion System for space transportation vehicles, while, at the same time, there is increasing concern about its continued affordability and rather involved reusability. Two broad approaches to addressing this overall launch cost problem consist in one, the further development of the rocket motor, and two, the use of airbreathing propulsion to the maximum extent possible as a complement to the limited use of a conventional rocket. In both cases, a single-stage-to-orbit (SSTO) vehicle is considered a desirable goal. However, neither the "all-rocket" nor the "all-airbreathing" approach seems realizable and workable in practice without appreciable advances in materials and manufacturing. An affordable system must be reusable with minimal refurbishing on-ground, and large mean time between overhauls, and thus with high margins in design. It has been suggested that one may use different engine cycles, some rocket and others airbreathing, in a combination over a flight trajectory, but this approach does not lead to a converged solution with thrust-to-mass, specific impulse, and other performance and operational characteristics that can be obtained in the different engines. The reason is this type of engine is simply a combination of different engines with no commonality of gas flowpath or components, and therefore tends to have the deficiencies of each of the combined engines. A further development in this approach is a truly combined cycle that incorporates a series of cycles for different modes of propulsion along a flight path with multiple use of a set of components and an essentially single gas flowpath through the engine. This integrated approach is based on realizing the benefits of both a rocket engine and airbreathing engine in various combinations by a systematic functional integration of components in an engine class usually referred to as a rocket-based combined cycle (RBCC) engine. RBCC engines exhibit a high potential for lowering the operating cost of launching payloads into orbit. Two sources of cost reductions can be identified. First, RBCC powered vehicles require only 20% takeoff thrust compared to conventional rockets, thereby lowering the thrust requirements and the replacement cost of the engines. Second, due to the higher structural and thermal margins achievable with RBCC engines coupled with a higher degree of subsystem redundance lower maintenance and operating cost are obtainable.

Siebenhaar, A.↗
NASA NTRS2008

Rapid Trajectory Optimization for the ARES I Launch Vehicle

A simplified ascent trajectory optimization procedure has been developed with application to NASA's proposed Ares I launch vehicle. In the interest of minimizing bending loads and ensuring safe separation of the first-stage solid rocket motor, the vehicle is con- strained to follow a gravity-turn trajectory. This reduces the design space to just two free parameters, the pitch rate after a short vertical rise phase to clear the launch pad, and initial launch azimuth. The pitch rate primarily controls the in-plane parameters (altitude, speed, flight path angle) of the trajectory whereas the launch azimuth primarily controls the out-of-plane portion (velocity heading.) Thus, the optimization can be mechanized as two one-dimensional searches that converge quickly and reliably. The method is compared with POST-optimized trajectories to verify its optimality.

Dukeman, Greg A.↗