Atmospheric Flight Phase Research Articles

Analytical trajectory solutions for atmospheric pull-up phase of interplanetary return flight

As a study on the flight of an interplanetary spacecraft back to Earth, the approximate analytical solutions for the atmospheric trajectory from reentry interface to pull-up point are derived. In future research, the solutions will be used as the theoretical tools for designing an analytical predictor-corrector guidance algorithm to control the pull-up flight. To derive the solutions, the highly nonlinear reentry dynamics is simplified first by planning the profile of drag-coefficient as an inversely proportional function of energy and using atmospheric density as a new state variable to replace altitude. Then, in order to decouple the dynamical equations, reference atmospheric density and reference flight-path angle are introduced to replace some real state variables while the differences between the reference and real variables are treated as small perturbations. Finally, a perturbation method is proposed to decompose the dynamics into two analytically solvable sub-systems. Thus the approximate analytical solutions are derived successfully. The simulation results demonstrate that the new solutions are more accurate than the existing solutions, because the new solutions give full consideration to the influence of Earth's curvature and rotation.

Evolution of Crew Escape System Configuration

The Crew Escape System (CES) is used for safe return of crew in case of any malfunctioning of a launch vehicle during ascent, specially in the atmospheric phase of flight. In the present case, as no active control is employed, the aerodynamic stability of CES is essential for safe return of crew in case of any emergency during ascent phase. While finalizing the shape of CES, besides stability, other aerodynamic parameters like acoustic, thermal, structural load, etc. need to be analysed. Based on detailed studies using CFD, a conical-shaped CES has been finalized. This article discusses the challenges involved in aerodynamic shaping of CES and different parameters studied while finalizing the configuration.

Mission Design, Preflight and Flight Performance and Observations for Pad Abort Test

Abort system is initiated to take the Crew Module (CM) away from the launch vehicle in case of an emergency at lift-off or at any point of time after launch for a mission with crew onboard. Crew Escape System (CES)-based abort is carried out from launch pad and during the atmospheric phase of ascent flight. The design and operation of CES play a crucial role in providing abort capability for escape from launch vehicle and return of the crew back to Earth during critical phase of ascent flight. CES motors are used to pull CM away from the launch vehicle during this mode of abort. Mission simulation and analysis is necessary for the design of CES-based abort mission and for its configuration. This article discusses the mission design, challenges faced during the design and strategies formulated towards the successful execution of Pad Abort Test.

Genetic Algorithm Based Parameter Tuning for Robust Control of Launch Vehicle in Atmospheric Flight

A hybrid approach to the design of the attitude control system for a launch vehicle (LV) in the atmospheric flight phase is proposed in this paper, where a structured <i xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">H</i> <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">∞</sub> controller is tuned using a genetic algorithm (GA). The <i xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">H</i> <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">∞</sub> synthesis relies on a classical architecture for the thrust vector control (TVC) system that features proportional-derivative loops and bending filters. Once a set of requirements on stability and robustness typical of industrial practice is specified, control design is carried out by parameterizing the <i xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">H</i> <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">∞</sub> weighting functions, and solving a two-layer max-min global optimization problem for the tuning parameters. The design methodology is applied to the model of a medium-size LV. The novel design is analyzed in off-nominal conditions taking into consideration model parameter scattering and wind disturbances. The results show that the automated design procedure allows to devise time-scheduled controllers providing adequate stability and performance, and appears as a viable and effective solution in order to reduce the burden of recurrent activities for controller tuning and validation conducted prior to each launch.

Development and Evaluation of AI System to Specify Favourable Launch Windows of Space Vehicles from the MST Radar Data

Weather being a complex Nondeterministic Polynomial time System, forecasting weather parameters has been a major challenge for so long and continues to be so. To address the pressing challenges, new prediction models are developed that vary in the prediction duration and accuracy. Predicting the wind speed, wind direction and turbulence is of significant interest to rocket industry as that would leverage them to time their launches. Launch vehicle trajectory would use the altitude based on mean wind biasing during atmospheric phase of flight to reduce significantly the effect of atmospheric winds on the structural loads experienced by the vehicle. Biasing the open loop pitch and yaw programs to the mean wind condition of launch month is known as monthly mean wind biasing. Atmospheric phase of flight refers to the wind profile up to 20-25 km altitude, the height resolution and accuracy is important. Diurnal variation on the day of launch is critical. Surface winds are measured at various tower heights up to 150 metres. While upper winds are measured with radiosonde mounted to Rawin balloons or with the expensive and difficult to procure Jimsphere or with MST radars. MST radar is far superior to the traditional radiosonde mounted to Rawin balloons in that the MST radar can be operated on a continuous basis to obtain measurements of the wind. MST Radar data is considered in the present study. Usually, numerical methods involving complex mathematical modelling else statistical estimates are employed for weather forecasting. In this article, an appropriate soft computing technique such as the Feed Forward based Back Propagation Neural Networks is proposed. Neural networks are good at fitting any nonlinear nature of function with several mathematical learning co-efficient variables. Analyses are carried out using MATLAB software for assessing the launch windows of space vehicles and its applications.

Finite Horizon Worst Case Analysis of Launch Vehicles

Abstract This paper presents a robust linear time varying (LTV) analysis of a space launcher. It evaluates the worst case aerodynamic loads during the atmospheric flight phase under wind turbulence. The results of the LTV analysis are evaluated against a Monte Carlo simulation of the respective industry-sized nonlinear launcher model over a set of allowable uncertainties and disturbances. The recent extension of the strict/ finite horizon bounded real lemma (BRL) for integral quadratic constraints is applied, covering the robustness analysis of finite horizon linear time varying (LTV) systems. Here, an equivalent formulation of the strict BRL in the form of a Riccati differential equality (RDE) is used. The RDEs direct dependence on the IQC multiplier leads to a nonlinear optimization problem, which can be solved efficiently. Using this, it is possible to explicitly respect the time varying dynamics of the space launcher in the worst case analysis. Furthermore, the incorporation of IQCs opens the analysis to a broader range of analyzable perturbations.

Minimum-Fuel Low-Earth-Orbit Aeroglide and Aerothrust Aeroassisted Orbital Transfer Subject to Heating Constraints

A numerical optimization study of minimum-fuel low-Earth-orbit aeroglide and aerothrust aeroassisted orbital transfer of a small spacecraft subject to constraints on heating rate and heating load with a required inclination change is considered. The aeroassisted orbital transfer is formulated as a three-phase optimal control problem consisting of an exoatmospheric deorbit phase, an atmospheric flight phase that may or may not include thrust, and a second exoatmospheric flight phase. The two different variations of the three-phase optimal control problem that arise from the trajectory design are solved using a high-accuracy adaptive Gaussian quadrature collocation method. The fuel consumption of both the aeroglide and aerothrust aeroassisted maneuvers are assessed as functions of the maximum allowable heating rate, the maximum allowable integrated heat load, the maximum lift-to-drag ratio of the vehicle, and the initial mass of the vehicle. Finally, the key features of the optimal aeroglide and aerothrust maneuvers are identified.

Robust Design of H∞Controller for a Launch Vehicle Autopilot against Disturbances

Atmospheric flight phase of a launch vehicle is utilized to evaluate the performance of an H-infinity controller in the presence of disturbances. Dynamics of the vehicle is linearly modeled using time-varying parameters. An operating point was found to design a robust command tracker using H-infinity control theory that guarantees a stable maneuver. At the end, the controller was employed on the launch vehicle to assess the capability of control design on the linearized aerospace vehicle. Experimental results illustrate the excellent performance of the H-infinity controller and accurate tracking implemented by the autopilot. Also the robustness of the entire system against disturbances is demonstrated to be acceptable.

Application of Closed Loop Optimal Guidance for a Constant Thrust Space Vehicle

This research presents derivation and implementation of the explicit guidance problem to steer a space vehicle into exo atmospheric phase of flight to develop three-dimensional optimal trajectory. The proposed guidance algorithm is in association with continuous powered flight of the space vehicle in ascent manoeuvre. Stability, accuracy and simplicity of this approach are the improved developments in comparison with the IGM approach. This algorithm uses the calculus of variation method for the two boundary-value injection problem to generate an optimal trajectory of space vehicle with online generation of steering command to inject to any desired orbit. Here the end conditions have been determined as the orbital height, inclination and eccentricity where the initial conditions are fixed. The simulation results are considered which shows the accuracy and simplicity of this method to reach to the desired orbit in minimum fuel.

Worst Case Analysis of a Launcher Vehicle Using Surrogate Models

Polynomial surrogate models of an industrial standard, closed-loop launcher model during an atmospheric flight phase are developed. A probabilistic collocation method is applied to derive surrogate models of varying complexity for wind perturbations occurring at different time instances during the atmospheric flight phase. These models are employed in optimisation-based worst-case analysis to significantly reduce the computational overhead while maintaining the accuracy of the computed worst-case uncertain parameter combinations. Comparisons of worst-case analysis results obtained using standard simulation models and their polynomial surrogate equivalents demonstrate significant potential of the proposed approach for radically reducing the computational burden associated with the industrial verification and validation process.

Robust safety margin assessment and constrained worst-case analysis of a launcher vehicle

The robustness of an industrial standard, closed-loop launcher model during an atmospheric flight phase is assessed in terms of robust safety margins in a multi-dimensional parameter space made up of various rigid and bending mode uncertainties. A single-axis launcher model with time-varying aerodynamic and thruster efficiency parameters, along with an H∞ controller and bending mode filter is considered. A constrained hybrid optimisation approach is adopted to determine the safety margins for different wind perturbations acting at different time instances of the flight. In addition, optimisation-based analysis is performed to determine the worst-case uncertain parameter combination for the launcher during the atmospheric flight phase. Probability constraints on the uncertain parameter combination satisfying the dispersion requirements are considered during the analysis.

Doppler effect in aeroacoustics

Doppler effect in aeroacoustics The prediction of acoustic fallout caused at the ground level by aircraft in flight or by launchers at take‐off sets the problem of the distortion of the emitted sound spectrum by the Doppler effect. Former studies applied to this problem use a geometrical formalism based on simplifying hypotheses such as homogeneous atmosphere, rectilinear trajectory and constant velocity of the sound source. Other studies consider the supersonic case using a model of spherical waves inside the Mach cone. These approaches are in fact not well adapted to the real cases, and the literature does not give many practical examples concerning this topic. In this study, we propose a new time approach of the Doppler effect that allows to avoid the complexity of the geometrical models and to take into account real atmosphere, curvilinear trajectories or accelerations of moving sound sources. This model allows to easily calculate the frequency shift and the change of acoustic level of the signal recorded at a given hearing point, in both cases of subsonic and supersonic sound sources. We present a first validation of this method applied to signals recorded in Kourou during the atmospheric phase of Flight 521 of Ariane 5.

Non-Stationary H∞ Control Law for Launcher with Bending Modes

This paper presents the development of an H∞ robust control law for the atmospheric flight phase of a future Ariane's type launcher. Some particular points of the development are the launcher control problem transcription into H∞ methodology framework, the take into account of system non-stationary features and the problem of finding a simplified on-board implementation of H∞ controllers. Satisfactory results were obtained in terms of a synthesis methodology definition, H∞ law performances compared to an LQG control law reference and implementation solutions with an LQG type structure for H∞ controllers.

Hybrid Near-Optimal Atmospheric Guidance for Aeroassisted Orbit Transfer

In this paper, a hybrid methodology is applied to the problem of determining a near-optimal control which approximately minimizes energy loss for a class of aeroassisted orbit transfer maneuvers. The result is an ap- proximate solution incorporating aspects of both an analytical zero-order solution from regular perturbation and a numerical solution from collocation. This approach is used to compute both open-loop and closed-loop near- optimal controls for a particular vehicle. Results from several approximations are quantitatively compared with an exact optimal solution. N orbit plane change is one of the most fuel-intensive ma- neuvers that can be required of an orbiting space vehicle. It has been shown that an aeroassisted orbit transfer consumes signifi- cantly less fuel than a similar maneuver executed with thrusters only. This advantage is only realizable, however, when optimal control laws are developed that take full advantage of the available aerody- namic forces. Because the dynamics governing this type of maneu- ver are highly nonlinear, an analytical solution to the optimal control problem is impossible. Furthermore, an exact numerical solution is too computationally demanding for real-time implementation on- board the orbiting vehicle. Approximation techniques that allow near-optimal guidance laws to be determined either analytically or with a minimum of computational effort are, therefore, highly de- sirable for this application. In this paper, a method is considered that combines elements of both approaches. An analytical approx- imation that holds for the atmospheric phase of flight is obtained from the zero-order solution to a regular perturbation problem. This solution is then augmented numerically by a technique borrowed from the method of collocation, in which the trajectory is divided into finite elements. In recent efforts, numerous approximate methods have been applied to the task of determining a near-optimal control for aeroassisted orbit transfer. These approximations have addressed the atmospheric portion of the maneuver, assuming that the effects of the orbital dynamics are small.13 Keplerian effects are treated by introducing an approximation of Loh's term. Some approxima- tions of this type have included assuming that Loh's term is con- stant or piecewise constant.2 A second alternative is to account for Keplerian effects through a regular perturbation expansion of the control solution.1 A related approach to the problem addresses both the orbital and atmospheric stages of the trajectory by invoking sin- gular perturbation theory and matched asymptotic expansions.4 It has been shown in Ref. 4 that the problem in question is, in fact, a singularly perturbed problem and that variations in Loh's term can be accounted for by properly matching a zero-order outer solution with the zero-order solution used in Ref. 1. That is, the zero-order solution of Ref. 1 serves as the inner solution in a matched asymp- totic expansion. The paper begins with a review of the dynamics of aeroas- sisted orbit transfer and the associated optimal control problem. A summary of the problem formulation as presented in Ref. 1 is included in the interests of clarity and completeness. The prob- lem definition is followed by a description of its analytical zero- order solution. An overview of the method of collocation is then presented as an entirely numerical methodology for this prob- lem. Next, the hybrid methodology incorporating elements of both solutions is described. This method was originally developed in Ref. 5 and applied to the problem of launch vehicle guidance. Finally, numerical results are presented for the maneuverable re- search re-entry vehicle (MRRV) which was featured in Ref. 3. These results are discussed in detail with recommendations for further analysis. Aeroassisted Orbit Transfer Problem The aeroassisted orbit transfer maneuver may be defined as an optimal control problem. This problem is characterized by highly nonlinear coupled differential constraints and is not amenable to an analytical solution. The equations may, however, be transformed and then used under certain simplifying assumptions to determine a near-optimal control which is valid for the atmospheric portion of the maneuver. In this section, the equations of motion which govern this type of maneuver are presented, and a scheme for modifying the equations is reviewed. The associated optimal con- trol problem is then defined, along with the general form of its solution.

Shuttle orbiter guidance system for the terminal flight phase

During the terminal atmospheric flight phase, the Space Shuttle orbiter is an unpowered glide vehicle. The guidance and flight control system during this phase must control energy, flight path, and heading to align the orbiter with the runway at an appropriate altitude and speed for an unpowered landing. This paper includes (1) a discussion of functional requirements, (2) definition of the automatic guidance and control system, (3) definition of the displays for pilot monitoring and control, and (4) a discussion of the dynamic performance of the system.

Shuttle Orbiter Guidance System for the Terminal Flight Phase

During the terminal atmospheric flight phase, the Shuttle orbiter is an unpowered glide vehicle. The guidance and flight control system during this phase must control energy, flight path, and heading to align the orbiter with the runway at an appropriate altitude and speed for an unpowered landing. The paper presents a technical description of the guidance and control system used for the terminal flight phase, which includes (1) a discussion of functional requirements. (2) definition of the automatic guidance and control system, (3) definition of the displays for pilot monitoring and control, and (4) a discussion of the dynamic performance of the system.