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This paper describes a control scheme used to quickly define a complicated trajectory for analysis of aircraft and spacecraft operations for which closed-form equations are not suitable. A “Trajectory Script” is a sequence of event-driven flight control commands that define a trajectory and “fly” the vehicle during simulation. A Trajectory Script can be developed in less than a minute, and can be changed and rerun in seconds. The method is implemented in the ROAST trajectory module of the RDS win aircraft design software, but could be applied to any time-dependent trajectory analysis. The paper describes the creation and operation of Trajectory Scripts, and discusses specific coded algorithms essential to implementing the concept in a time-domain trajectory simulation program.

This paper proposes novel orbits in the Circular Restricted Three-Body Problem (CRTBP) of the Mars-Phobos system. We have found different types of quasiperiodic orbits with long-term stability that can be connected into the invariant manifolds that transport the spacecraft directly into a L1 halo orbit. We analyze the feasibility of several transfer trajectories connecting different sizes of halo orbits around L1 and L2. We further analyze the station-keeping maneuvers for the halo orbits around the collinear libration points allowing the spacecraft to stay for several days and increase the science performed during proximity operations in the vicinity of Phobos.

This paper presents results from a contingency trajectory analysis performed for the Lunar Atmosphere & Dust Environment Explorer (LADEE) mission in the event of a missed lunar-orbit insertion (LOI) maneuver by the LADEE spacecraft. The effects of varying solar perturbations in the vicinity of the weak stability boundary (WSB) in the Sun-Earth system on the trajectory design are analyzed and discussed. It is shown that geocentric recovery trajectory options existed for the LADEE spacecraft, depending on the spacecraft's recovery time to perform an Earth escape-prevention maneuver after the hypothetical LOI maneuver failure and subsequent path traveled through the Sun-Earth WSB. If Earth-escape occurred, a heliocentric recovery option existed, but with reduced science capacapability for the spacecraft in an eccentric, not circular near-equatorial retrograde lunar orbit.

In this paper trajectory analysis for the aerial refueling basket for a modern fighter aircraft is presented. The under study aircraft has a probe and drogue arrangement for aerial refueling. Thus the aircraft has to be carefully positioned behind the extended refueling basket is such a way that positive engagement of refueling basket and probe takes place accounting for the complex flow field around the aircraft as it approaches the refueling basket. In the current study an initial location of the refueling basket has been selected such that, the aircraft flow field have minimal aerodynamic effects on the basket at this location and, then the basket is allowed to move towards the aircraft using Six Degree of Freedom Solver with appropriate boundary and limiting conditions. The initial location of the basket is iteratively improved until engagement of refueling probe and basket is achieved. Further optimization of initial location and some worst case scenarios have also been simulated. The results indicate significant effect of aircraft aerodynamics on basket trajectory. An optimum initial basket location that results in positive engagement with the refueling probe has been determined.

Functionality has recently been added to Boeing’s Trajectory Analysis and Modeling Environment (TAME) to model the Air Traffic Controller (ATC) use of a new tool, Relative Position Indicator (RPI), developed by MITRE CAASD. This tool is currently being demonstrated and tested at select FAA TRACON facilities within the NAS. The RPI is a visual aid that can be used by controllers to quickly assess diverse merging situations within the TRACON. The RPI is currently being postulated as a means to enable Optimal Profile Descent (OPD) operations in the TRACON while aircraft fly on published RNAV procedures, using speed control to manage traffic merging. While the RPI tool projects aircraft “ghost” positions using distance rather than estimated time of arrival (ETA) to the merge point, the controllers compensate for this effect in their speed clearances using knowledge of wind, altitude, and speed schedules for the paths involved, and it is this functionality that is modeled in TAME. This paper describes the model of controllers using the RPI in TAME, called the RPI Emulator, and a set of trade studies designed to evaluate the use of RPI in a complex TRACON operation. The results test the effectiveness of speed control as a mechanism to manage merging traffic on RNAV procedures in the TRACON, and the relationship between this separation method and the design of the RNAV procedures.

The conceptual design activities of the development of the SCRAMSPACE I scramjet-powered free-flight experiment are described in this paper. The objectives of this flight are first described together with the definition of the primary, secondary and tertiary experiments. The Scramjet configuration studied is first discussed together with the rocket motor system selected for this flight. The different flight sequences are then explained, highlighting the SCRAMSPACE I free-flyer separation and re-orientation procedures. A trajectory analysis and the derived thermo-structural analysis are presented, discussing the alignment of the predicted performance with the mission scientific requirements. The global system architecture and instrumentation of the vehicle are then explained. The conclusions of these Phase A technical activities are that the presented mission concept is in line with the SCRAMSPACE I scientific requirements and no major technical difficulties have been identified.

The use of electric propulsion for near-Earth application has become routine and its capabilities continues to grow. However, there is limited experience with electric propulsion for planetary science missions. Low-thrust trajectory analysis is considered an art more than science with independent analyses producing inconsistent results. Planetary missions often have characteristics counterintuitive to standard ballistic trajectory design. A standard application of subsystem margins with the use of electric propulsion may not be appropriate due to margin associated interdependencies. An introduction to primary electric prolusion thrusters, mission design characteristics associated with low-thrust implementation and application are presented.

Various lunar descent trajectories were analyzed that include the optimization of the Apollo constrained mission trajectory, a fully optimized minimum energy trajectory, and a optimal, constrained trajectory using current instrumentation technology. Trade studies were conducted to determine the impacts of mission assumptions, pilot in the loop/automated flight demands, and additional constraints for the present recurring missions to the same outpost landing site. For mission design at this conceptual phase of the program, the Apollo pre-mission planning was applied to account for known contingencies (hardware, instrumentation known uncertainties) and unknown unknowns. The mission Delta-V’s are presented in a risk form of conservative, nominal, and optimistic range where 90 percent of Delta-V is derived by trajectory analysis and the other 10 percent was derived from a qualitative analysis from Apollo 11 pre-mission planning. The recommendations for the Delta Vs are the following: conservative (Apollo derived) (2262 m/s), nominal (2053 m/s), and optimistic (1799 m/s). Because of the qualitative nature of the results, the degree of autonomy assumed, additional safety considerations for a lunar outpost, and the impact of advanced instrumentation, more in-depth analyses are required to refine the present recommendations. I. Nomenclature

Six degree-of-freedom (DOF) launch vehicle trajectories are designed to follow an optimized 3-DOF reference trajectory. A vehicle has a finite amount of control power that it can allocate to performing maneuvers. Therefore, the 3-DOF trajectory must be designed to refrain from using 100% of the allowable control capability to perform maneuvers, saving control power for handling off-nominal conditions, wind gusts and other perturbations. During the Ares I trajectory analysis, two maneuvers were found to be hard for the control system to implement; a roll maneuver prior to the gravity turn and an angle of attack maneuver immediately after the J-2X engine start-up. It was decided to develop an approach for creating smooth maneuvers in the optimized reference trajectories that accounts for the thrust available from the engines. A feature of this method is that no additional angular velocity in the direction of the maneuver has been added to the vehicle after the maneuver completion. This paper discusses the equations behind these new maneuvers and their implementation into the Ares I trajectory design cycle. Also discussed is a possible extension to adjusting closed-loop guidance.

This viewgraph presentation examines flight hardware elements of the Constellation Program (CxP) and the utilization of the Crew Exploration Vehicle (CEV), Evolvable Expendable Launch Vehicles (EELVs) and Ares launch vehicles for NEO missions.

NASA s highly successful Hyper-X program demonstrated numerous hypersonic air-breathing vehicle related technologies including scramjet performance, advanced materials and hot structures, GN&C, and integrated vehicle performance resulting in, for the first time ever, acceleration of a vehicle powered by a scramjet engine. The Systems Analysis and Concepts Directorate (SACD) at NASA s Langley Research Center played a major role in the integrated team providing critical support, analysis, and leadership to the Hyper-X Program throughout the program s entire life and were key to its ultimate success. Engineers in SACD s Vehicle Analysis Branch (VAB) were involved in all stages and aspects of the program, from conceptual design prior to contract award, through preliminary design and hardware development, and in to, during, and after each of the three flights. Working closely with other engineers at Langley and Dryden, as well as industry partners, roughly 20 members of SACD were involved throughout the evolution of the Hyper-X program in nearly all disciplines, including lead roles in several areas. Engineers from VAB led the aerodynamic database development, the propulsion database development, and the stage separation analysis and database development effort. Others played major roles in structures, aerothermal, GN&C, trajectory analysis and flight simulation, as well as providing CFD support for aerodynamic, propulsion, and aerothermal analysis.

On the morning of January 15, 2006, the Stardust capsule successfully landed at the Utah Test and Training range in northwest Utah returning cometary samples from the comet Wild-2. An overview of the entry, descent, and landing (EDL) trajectory analysis that was performed for targeting during the Stardust Mission Navigation Operations Phase upon final approach to Earth is described. In addition, how the predicted landing location and the resulting overall 99 percentile landing footprint ellipse obtained from a Monte Carlo analysis changed over the final days and hours prior to entry is also presented. The navigation and EDL operations effort accurately delivered the entry capsule to the desired landing site. The final landing location was 8.1 km from the target, which was well within the allowable landing area.

The Irvin Aerospace (Irvin) trajectory and parachute inflation tool DCLDYN (Decelerator Dynamics) has been in existence for many years, with algorithmic portions beginning in the later stages of the Apollo program. The ensuing years, particularly the past ten, have provided a number of added simulation capabilities and upgrades, building on these basic algorithms. Some of these capabilities include

This paper is part of a sequence of three documenting the work completed at the NASA Marshall Space Flight Center in the area of the defense of Earth from incoming Near Earth Objects (NEO). The work found herein can be found fully docum ented in reference 1 . I. Outbound Propulsion For both the interception and rendezvous techniques, neither fragmentation nor deflection can take place until the necessary system hardware is transported out to the approaching NEO. Some type of outbound pro pulsion system is required to accomplish this. Several propulsion systems are considered in this study; they were selected for their ability to meet the mission requirements, their level of technological maturity and development status.

This paper describes the engineering of several vehicles designed for a crewed mission to the Jovian satellite Callisto. Each subsystem is discussed in detail. Mission and trajectory analysis for each mission concept is described. Crew support components are also described. Vehicles were developed using both fission powered magneto plasma dynamic (MPD) thrusters and magnetized target fusion (MTF) propulsion systems. Conclusions were drawn regarding the usefulness of these propulsion systems for crewed exploration of the outer solar system.

A new methodology has been implemented to provide a higher level of usability for NASA’s Scientific Balloon Analysis Model, SINBAD, while also enhancing the interaction among the community of users. The methodology sequesters the thermal trajectory analysis inputs into the following seven categories: the balloon, the atmosphere, the gas, the film, the auto-ballaster, vents & valves, and the launch. By implementing this methodology in a graphical user interface, the user can test an existing balloon in different atmospheres, with different film properties, etc., mixing and matching inputs quickly. Using this database sharing system, users can also export analyses to files that may be shared via e-mail and remerged by remote users, which increases the depth of communication between analysts. The problems associated with designing such a system include defining the categories and determining how the simulator input is distributed into those categories. These problems, along with other design and implementation issues are discussed.

Two primary simulations have been developed and are being updated for the Mars Smart Lander Entry, Descent, and Landing (EDL). The high fidelity engineering end-to-end EDL simulation that is based on NASA Langley's Program to Optimize Simulated Trajectories (POST) and the end-to-end real-time, hardware-in-the-loop simulation testbed, which is based on NASA JPL's (Jet Propulsion Laboratory) Dynamics Simulator for Entry, Descent and Surface landing (DSENDS). This paper presents the status of these Mars Smart Lander EDL end-to-end simulations at this time. Various models, capabilities, as well as validation and verification for these simulations are discussed.

An integrated trajectory analysis, aerothermal analysis and Thermal Protection System (TPS) sizing capability has been developed. The method employs the Traj trajectory tool, the GASP computational fluid dynamics solver and the FIAT thermal protection system sizing routine. Results are presented for a baseline configuration of the Mars Smart Lander (MSL) entry vehicle. Design issues and TPS selection and weights are revealed. The capability represents recent efforts at the NASA Ames Research Center to support integrated TPS design.

Fully reusable two-stage-to-orbit vehicle designs that incorporate ObranchingO trajectories during their ascent are of current interest in the advanced launch vehicle design community. Unlike expendable vehicle designs, the booster of a reusable system must fly to a designated landing site after staging. Therefore, both the booster return branch and the orbital upper stage branch along with the lower ascent trajectory are of interest after the staging point and must be simultaneously optimized in order to achieve an overall system objective. Current and notable designs in this class include the U. S. Air Force Space Operations Vehicle designs with their Opop-upO trajectories, the Kelly Astroliner, the Kistler K-1, the two-stage-to-orbit vehicle Stargazer, and NASAOs proposed liquid flyback booster designs (Space Shuttle booster replacement). The solution to this problem using an industrystandard trajectory optimization code (POST) typically requires at least two separate computer jobs N one for the orbital branch, from the ground to orbit, and one for the flyback branch, from the staging point to the landing site. These jobs are coupled and their data requirements are interdependent. These requirements must be taken into consideration when optimizing the entire trajectory. This paper analyzes the results of branching trajectory optimization for the Kistler K-1 launch vehicle with respect to computational efficiency

The scope of the present investigation includes realistic trajectory simulations for a biconic aerocapture vehicle Copyright© 1998 by North Carolina State University. Published by the AIAA, Inc. with permission. Graduate Research Assistant, Student AIAA. ^ Professor, Department of Mechanical and Aerospace Engineering, and Director, Mars Mission Research Center. Fellow AIAA. with an L/D of 0.78, determination of its entry corridor, design of its nominal entry trajectory, evaluation of it's tolerance to atmospheric density variation, which can be possibly encountered during aerocapture, and the change in propellant usage resulting from these dispersions. To make the simulation more realistic, minimum €3 mission opportunities for the years 2001, 2003 and 2005 are chosen for consideration and their results are compared. Arrival dates, entry velocities and geographical entry locations for these mission opportunities' are used as entry conditions. In addition, comparative trajectory analysis between a biconic vehicle( Z/£>=0.78 ) and a Viking type vehicle (L/D=0.18) are carried out for the 2005 mission. The 2005 opportunity is a likely candidate for a Mars Sample Return Mission which could benefit significantly from aerocapture.

A recently developed algorithm for nonlinear system performance analysis has been applied to an F16 aircraft to begin evaluating the suitability of the method for aerospace problems. The algorithm has a potential to be much more efficient than the current methods in performance analysis for aircraft. This paper is the initial step in evaluating this potential.

In this paper, a hybrid methodology is used to determine optimal open loop controls for the atmospheric portion of the aeroassisted plane change problem. The method is hybrid in the sense that it combines the features of numerical collocation with the analytically tractable portions of the problem which result when the two-point boundary value problem is cast in the form of a regular perturbation problem. Various levels of approximation are introduced by eliminating particular collocation parameters and their effect upon problem complexity and required number of nodes is discussed. The results include plane changes of 10, 20, and 30 degrees for a given vehicle.

The methods to be described in this paper were developed during the analysis of trajectory data for the G-12E family of airdrop systems. They formed the basis for determining the minimum altitude above ground level AGL) at which aircraft may deliver cargo using L -1 2E parachutes without the cargo suffering impact damage. This was motivated by the U.S. Air Force's desire to minimize the risk of hostile fire by flying as low to the ground as possible during airdrop operations. The application of these methods are not limited to the "minimum altitude" problem; they can also be used to determine the quality of airdrop systems and components.

The conceptual design of a rocket-powered, two-stage fully reusable launch vehicle has been performed as a part of NASA's Advanced Manned Launch System (AMLS) study. This paper summarizes the structural design and loads analysis of this two-stage fully reusable launch vehicle. The method used to determine the structural weights consists of generating a finite-element model for each vehicle, selecting a set of critical loading conditions, determining the loads on the model caused by those conditions, determining the model response and changing the sizes of individual elements to obtain a safe structure. The integrated loads on the two-stage vehicle were obtained from a three-degrees-of-freedom trajectory analysis.

The National Launch System (NLS) is a joint concept, between DoD and NASA, for building a family of new launch vehicles. Two of the many choices to be made are the trajectory shaping methods and the onboard guidance scheme. This paper presents results from some ongoing studies to address these issues. First, potential gains from new guidance concepts are listed. Next the paper gives a list of possible discriminators between different guidance schemes, lists a number of potential guidance schemes, and explains two in some detail. A reference scheme is tested to determine its performance versus the discriminators. Finally, results from some special studies using the reference guidance scheme are given, including the effects of closed-loop guidance initiation time, time of enforcement of sideslip, vehicle roll for engine out, time and location of an engine out, use of load relief control, and use of day of launch wind biasing.

Approximately one hour prior to the September 1991 launch of the Space Shuttle Discovery on STS-48, new guidance commands were uplinked to the onboard computers, allowing the vehicle to fly safely through unusually strong upper atmosphere winds. The capability to update the vehicle guidance commands known as Day-of-Launch I-load Update (DOLILU), had been developed by a NASA/industry team, and was certified for flight use in August 1991 following an extensive testing regime. The DOLILU capability presents several benefits to the Space Shuttle Program. Not only does it reduce the possibility of a launch scrub due to unexpected wind profiles at high altitude, it is also a major first step toward standardization of the ascent flight profile, and which will reduce the cost of flight software verification commit-to-flight trajectory analysis, and flight crew training.

A two-phase experiment is proposed which utilizes the Shuttle Orbiter and its unique series of repeated entries into the earth's atmosphere as an airborne in situ aerodynamic testing laboratory. The objective of the experiment is to determine static aerodynamic force coefficients, first of the orbiter, and later of various entry configurations throughout the high speed flight regime, including the transition from free molecule to continuum fluid flow. The objective will be accomplished through analysis of inflight measurements from both shuttle-borne and shuttle-launched instrumented packages. Results are presented to demonstrate the feasibility of such an experiment.