PAYLOAD OPTIMIZATION FOR POWER-LIMITED VEHICLES
Optimization of propulsion system and trajectory parameters of a power-limited space vehicle to achieve maximum payload with application to earth-mars rendezvous missions
LOW THRUST PROPULSION↗
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Results for “vehicle trajectory”
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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Trim conditions of mated vehicles
Program was designed to generate data from effects of configuration and trajectory on 747/space shuttle vehicle. System could be adapted to any similar vehicle arrangement. For given coupled vehicle configuration, program will define rate-of-climb ceiling, corresponding trim attitude, and 747 surface deflection at any attainable altitude.
Nobles, C. M.↗
Analytical solutions to constrained hypersonic flight trajectories
The flight trajectory of aerospace vehicles subject to a class of path constraints is considered. The constrained dynamics is shown to be a natural two-time-scale system. Asymptotic analytical solutions are obtained. Problems of trajectory optimization and guidance can be dramatically simplified with these solutions. Applications in trajectory design for an aerospace plane strongly support the theoretical development.
Lu, Ping↗
Analytical solutions to constrained hypersonic flight trajectories
The flight trajectory of aerospace vehicles subject to a class of path constraints is considered. The constrained dynamics is shown to be a natural two-time-scale system. Asymptotic analytical solutions are obtained. Problems of trajectory optimization and guidance can be dramatically simplified with these solutions. Applications in trajectory design for an aerospace plane strongly support the theoretical development.
Lu, Ping↗
Optimal return-to-launchsite abort trajectories for an HL-20 Personnel Launch System vehicle
The Personnel Launch System (PLS) being studied by NASA is a system to complement the Space Shuttle and provide alternative access to space. The PLS consists of a manned spacecraft launched by an expendable launch vehicle (ELV). A candidate for the manned spacecraft is the HL-20 lifting body. In the event of an ELV malfunction during the initial portion of the ascent trajectory, the HL-20 will separate from the rocket and perform an unpowered return-to-launchsite (RTLS) abort. This paper describes an investigation of the RTLS abort scenario using optimal control theory. The objective of the abort trajectory is to maximize final altitude at a point near the runway. The assumption is then made that there exists a control history to steer the vehicle to any final altitude lower than the final optimal altitude. With this selection of cost function, and with this assumption, the feasibility of an RTLS abort at different times along the ascent trajectory can be determined. The method of differential inclusions, which allows the determination of optimal states and eliminates the need for determining the optimal controls, is used to determine the optimal trajectories.
Dutton, Kevin E.↗
The ballistic Mars hopper - An alternative Mars mobility concept
The ballistic Mars hopper is proposed as an alternative mobility concept for unmanned exploration of the Martian surface. In the concept, oxygen and carbon monoxide produced from the Martian atmosphere are used as propellants in a rocket propulsion system for an unmanned vehicle on suborbital trajectories between landing sites separated by distances of up to 1000 km. This mobility concept is seen as uniquely capable of allowing both intensive and extensive exploration of the planet using only a single landed vehicle of mass approximately 2000 kg. The technical challenges associated with in-situ propellant production on the surface of Mars are reviewed. A rocket propulsion subsystem capable of using oxygen and carbon monoxide as propellants is described. Finally, results of mission analysis and a hopper landing hazard simulation are reported. It is concluded that an attractive Mars hopper can be developed based on relatively near-term technology.
Sercel, Joel C.↗
Scientific objectives <of the solar probe<
Experiment designs, instrumentation, and trajectory for solar probe vehicle
SOLAR PROBE↗
Optimization and evaluation of main liquid rocket propulsion systems for advanced earth-to-orbit shuttles
A computerized preliminary design system is used to evaluate potential main liquid-rocket propulsion systems for advanced technology winged single-stage-to-orbit launch vehicles. Evaluated are trade-offs between ascent flight trajectory performance and flight vehicle sizing driven by engine mass and propellant requirements. Numerous mission, flight, and vehicle-related requirements and constraints are satisfied in the design process. With the design system, five dual-mode propulsion system concepts are compared to a baseline hydrogen and oxygen system in terms of the changes in vehicle dry mass and gross mass.
Wilhite, A. W.↗
Small Launch Vehicle Trade Space Definition: Development of a Zero Level Mass Estimation Tool with Trajectory Validation
Recent high level interest in the capability of small launch vehicles has placed significant demand on determining the trade space these vehicles occupy. This has led to the development of a zero level analysis tool that can quickly determine the minimum expected vehicle gross liftoff weight (GLOW) in terms of vehicle stage specific impulse (Isp) and propellant mass fraction (pmf) for any given payload value. Utilizing an extensive background in Earth to orbit trajectory experience a total necessary delta v the vehicle must achieve can be estimated including relevant loss terms. This foresight into expected losses allows for more specific assumptions relating to the initial estimates of thrust to weight values for each stage. This tool was further validated against a trajectory model, in this case the Program to Optimize Simulated Trajectories (POST), to determine if the initial sizing delta v was adequate to meet payload expectations. Presented here is a description of how the tool is setup and the approach the analyst must take when using the tool. Also, expected outputs which are dependent on the type of small launch vehicle being sized will be displayed. The method of validation will be discussed as well as where the sizing tool fits into the vehicle design process.
Waters, Eric D.↗
Optimal time trajectory and coordination for connected and automated vehicles
Not Available
Malikopoulos, Andreas A.↗
Rendezvous Integration Complexities of NASA Human Flight Vehicles
Propellant-optimal trajectories, relative sensors and navigation, and docking/capture mechanisms are rendezvous disciplines that receive much attention in the technical literature. However, other areas must be considered. These include absolute navigation, maneuver targeting, attitude control, power generation, software development and verification, redundancy management, thermal control, avionics integration, robotics, communications, lighting, human factors, crew timeline, procedure development, orbital debris risk mitigation, structures, plume impingement, logistics, and in some cases extravehicular activity. While current and future spaceflight programs will introduce new technologies and operations concepts, the complexity of integrating multiple systems on multiple spacecraft will remain. The systems integration task may become more difficult as increasingly complex software is used to meet current and future automation, autonomy, and robotic operation requirements.
Brazzel, Jack P.↗
Mars Sample Retrieval Lander Thermal Protection System Design
The Mars Sample Return (MSR) Sample Retrieval Lander (SRL) was slated to launch in 2031 and enter Mars’ atmosphere after a 15-month cruise. After several iterations of vehicle design and trajectories, the latest mission architecture included a 4.72-meter diameter entry capsule entering the Martian atmosphere at 8 km/s, marking a 3.5x increase in kinetic energy when entering the atmosphere in comparison to Mars 2020. Designed to be the heaviest and fastest vehicle to enter Mars’ atmosphere, SRL required slim mass margins, and as such, the thermal protection system (TPS) design methodology needed to be tailored appropriately. The entry capsule aeroshell is composed of a heatshield, backshell, parachute support cone, and parachute lid, as shown in Fig. 1.
Hannah Alpert↗
Prospects for utilization of air liquefaction and enrichment system (ALES) propulsion in fully reusable launch vehicles
A concept is shown for a fully reusable, earth to orbit launch vehicle with horizontal takeoff and landing, employing an air-turborocket for low speed and a rocket for high speed acceleration, both using LH2 fuel. The turborocket employs a modified liquid air cycle to supply the oxidizer. The rocket uses 90 percent pure LOX that is collected from the atmosphere, separated, and stored during operation of the turborocket from about Mach 2 to Mach 5 or 6. The takeoff weight and the thrust required at takeoff are markedly reduced by collecting the rocket oxidizer in-flight. The paper shows an approach and the corresponding technology needs for using ALES propulsion in a SSTO vehicle. Reducing the trajectory altitude at the end of collection reduces the wing area and increases payload. The use of state-of-the-art materials, such as graphite polyimide, is critical to meet the structure weight objective for SSTO. Configurations that utilize 'waverider' aerodynamics show great promise to reduce the vehicle weight.
Bond, W. H.↗
Adaptive Critic Neural Network-Based Terminal Area Energy Management and Approach and Landing Guidance
Reusable Launch Vehicles (RLVs) have different mission requirements than the Space Shuttle, which is used for benchmark guidance design. Therefore, alternative Terminal Area Energy Management (TAEM) and Approach and Landing (A/L) Guidance schemes can be examined in the interest of cost reduction. A neural network based solution for a finite horizon trajectory optimization problem is presented in this paper. In this approach the optimal trajectory of the vehicle is produced by adaptive critic based neural networks, which were trained off-line to maintain a gradual glideslope.
Grantham, Katie↗
OPTIMUM EARTH-TO-MARS ROUNDTRIP TRAJECTORIES UTILIZING A LOW-THRUST POWER-LIMITED PROPULSION SYSTEM
Payload capabilities of vehicle containing low- thrust power-limited propulsion system for earth- to-mars roundtrip mission
PAYLOAD↗
Calculation of free-fall trajectories using numerical optimization methods.
An important problem in space flight is the calculation of trajectories for nonthrusting vehicles between fixed points in a given time. A new procedure based on Hamilton's principle for solving such two-point boundary-value problems is presented. It employs numerical optimization methods to perform the extremization required by Hamilton's principle. This procedure is applied to the calculation of an Earth-Moon trajectory. The results show that the initial guesses required to obtain an iteration procedure which converges are not critical and that convergence can be obtained to any predetermined degree of accuracy.
Hull, D. G.↗
USE OF CALCULUS OF VARIATIONS METHODS FOR TRAJECTORY OPTIMIZATION AND ADVANCED GUIDANCE CONCEPTS
Calculus of variation methods used for trajectory optimization and guidance of rocket propelled vehicles
TRAJECTORY CONTROL↗
Development of Constraint Force Equation Methodology for Application to Multi-Body Dynamics Including Launch Vehicle Stage Seperation
The objective of this report is to develop and implement a physics based method for analysis and simulation of multi-body dynamics including launch vehicle stage separation. The constraint force equation (CFE) methodology discussed in this report provides such a framework for modeling constraint forces and moments acting at joints when the vehicles are still connected. Several stand-alone test cases involving various types of joints were developed to validate the CFE methodology. The results were compared with ADAMS(Registered Trademark) and Autolev, two different industry standard benchmark codes for multi-body dynamic analysis and simulations. However, these two codes are not designed for aerospace flight trajectory simulations. After this validation exercise, the CFE algorithm was implemented in Program to Optimize Simulated Trajectories II (POST2) to provide a capability to simulate end-to-end trajectories of launch vehicles including stage separation. The POST2/CFE methodology was applied to the STS-1 Space Shuttle solid rocket booster (SRB) separation and Hyper-X Research Vehicle (HXRV) separation from the Pegasus booster as a further test and validation for its application to launch vehicle stage separation problems. Finally, to demonstrate end-to-end simulation capability, POST2/CFE was applied to the ascent, orbit insertion, and booster return of a reusable two-stage-to-orbit (TSTO) vehicle concept. With these validation exercises, POST2/CFE software can be used for performing conceptual level end-to-end simulations, including launch vehicle stage separation, for problems similar to those discussed in this report.
Pamadi, Bandu N.↗