Abort Trajectories Research Articles

Design and analysis of direct abort trajectory for a crewed lunar mission based on the quasi-lambert problem

A direct design method for direct abort trajectory based on the quasi-lambert problem is proposed. The re-entry constraint is analytically addressed via a quasi-lambert problem instead of variables iteration and optimization method in the past. Then a new parameter is introduced to determine the plane. The exact solution is determined by differential correction based on the initial solution. Numerical simulation is carried out to test the validity and efficiency of the method. The comparison with STK verifies its reliability. Additionally, the characteristics of direct abort trajectory and their application in the option strategy for different abort ways are studied. The flight time and abort point affect the abort trajectories’ velocity increment and the latitude of the landing area.

Parameterized design of abort trajectories with a lunar flyby for a crewed mission

This article proposes an analytical preliminary design method for abort trajectory with a lunar flyby and the trajectory ergodic representation based on parameterization, as well as their application in studying the abort trajectories’ characteristics. First, by introducing two parameters, the perilune altitude and the perilune transfer time, the preliminary solution is developed parametrically using pseudostate theory to connect two special Lambert processes. The fast convergence of the improved differential correction under the high-fidelity gravity field verifies the accuracy of the proposed preliminary design method. Next, the two parameters are regarded as a set of variables to define an abort trajectory. Based on this, an ergodic representation of the abort trajectory with a certain abort time is yielded using the proposed design method, which offers a global view of the feasible abort trajectories. In addition, the influences of some factors on the performance of these abort trajectories are investigated based on the proposed design method and ergodic representation. The simulation results show that the abort time and the transfer time of the nominal trajectory significantly affect the ergodic representation and characteristics of the abort trajectory.

Characteristic analysis and design of near moon abort trajectory for manned lunar landing mission

The safety of astronauts would be severely threatened if the lunar-landing spacecraft were under an emergency during the near moon phase of flight, which was far from the Earth. For the problem of mission abort caused by the main engine (service propulsion system, SPS) failure during lunar orbit insertion, firstly, the family of trajectories resulted from SPS premature shutdown and corresponding abort trajectories were analyzed; then an algorithm that can be applied to the near moon abort trajectories was proposed using patched-conic technique. The characteristics of the abort trajectory, such as energy consumption and return time of flight, were analyzed and presented. Finally, simulation examples were given to demonstrate various cases of near moon SPS failure. The results of the simulation have validated the approach proposed.

Generation of Launch Vehicle Abort Trajectories Using a Hybrid Optimization Method

Modifications and extensions to a recently developed hybrid numerical–analytical method for optimization of future launch vehicles are presented, to allow for the real time generation of abort trajectories. A set of terminal constraints is formulated that targets an entry condition for the abort case and allows the vehicle to attach tangentially to a nominal entry path. The trajectories are optimal in the sense of maximizing terminal energy. In addition, the algorithm has been extended to include path constraints on both angle of attack and q ·α. A second issue addressed concerns the possibility of singular arcs, both for nominal ascent and aborts. It is proven that such arcs are not feasible, both within the atmosphere and in vacuum, when the control variables consist of body attitude and throttle. Finally, numerical results for different abort scenarios using a single-stage vehicle model illustrate both the capabilities and some of the deficiencies of the approach.

New Nonlinear Control Technique for Ascent Phase of Reusable Launch Vehicles

Modifications and extensions to a recently developed hybrid numerical-analytical method for optimization of future launch vehicles are presented, to allow for the real time generation of abort trajectories. A set of terminal constraints is formulated that targets an entry condition for the abort case and allows the vehicle to attach tangentially to a nominal entry path. The trajectories are optimal in the sense of maximizing terminal energy. In addition, the algorithm has been extended to include path constraints on both angle of attack and q . α. A second issue addressed concerns the possibility of singular arcs, both for nominal ascent and aborts. It is proven that such arcs are not feasible, both within the atmosphere and in vacuum, when the control variables consist of body attitude and throttle. Finally, numerical results for different abort scenarios using a single-stage vehicle model illustrate both the capabilities and some of the deficiencies of the approach.

Metallic Thermal Protection System Requirements, Environments, and Integrated Concepts

Achieving the ultimate goal of an economically viable reusable launch vehicle will eventually require developing Federal Aviation Regulation-type performance-based requirements and certification by the Federal Aviation Administration, as is currently done for commercial transports. Because the necessary requirements do not currently exist, there is no verifiable and traceable link between thermal protection system design implementation and resulting performance, safety, and cost. An initial attempt has been made to outline a set of performance-based thermal protection system design requirements. Critical requirements that will have a profound effect on the economic viability of a reusable launch vehicle, such as those for ground hail strike, lightning strike, bird strike, rain/rain erosion, and on-orbit dehris/micrometeoroid hypervelocity impact have been proposed. In addition to design requirements, the importance of both compiling a comprehensive loads envelope and deriving time- and location-consistent loads for thermal protection system design and sizing is addressed. Including ascent abort trajectories as limit-load cases and on-orbit debris/micrometeoroid hypervelocity impact as one of the discrete-source-damage cases is Imperative because of their significant impact on thermal protection system design and resulting performance, reliability, and operability. General features of a suite of integrated airframe concepts is summarized, and the specific details of a metallic thermal protection system concept having design flexibility that enables weight and operability to be traded and balanced is described.

Ablation analysis of the Shuttle Orbiter oxidation protected reinforced carbon-carbon

Reusable, oxidation-protected reinforced carbon-carbon has been successfully flown on all Shuttle Orbiter flights. Thermal testing of the silicon carbide-coated, reinforced carbon-carbon to determine its oxidation characteristics has been performed in convective (plasma Arc-Jet) heating facilities. Surface sealant mass loss was characterized as a function of temperature and pressure. High-temperature testing was performed to develop coating recession correlations for predicting performance at the over-temperature flight conditions associated with abort trajectories. Methods for using these test data to establish multimission reuse (i.e., mission life) and single mission limits are presented.