Launch Vehicle Guidance Research Articles

3D Profile Reconstruction and Guidance for the Terminal Area Energy Management Phase of an Unpowered RLV with Aerosurface Failure

AbstractThis paper focuses on the guidance of an unpowered reusable launch vehicle (RLV) with aerosurface failure in the terminal area energy management (TAEM) phase. The aerosurface failure will l...

Numerical and experimental comparative study of Aluminum and hybrid mounting interfaces of launch vehicle avionics for weight reduction

Weight is a paramount parameter for launch vehicle applications. Avionics are used for command, control and guidance of launch vehicle. Interface brackets are used for mounting of these avionic packages to the launch vehicle chassis. Dynamic loads and vibration responses dominate the design criteria for these interface brackets. The weight sensitivity on range is more prominent for the upper stages of the launch vehicle. To improve the performance of launch vehicle, employment of light weight hybrid (Aluminum honeycomb panel made up of honeycomb core with face sheets at top and bottom) bracket instead of comparatively heavy conventional all metallic aluminum alloy bracket, for a particular case of mounting two of the launch vehicle avionic packages is explored in this study. Comparable vibration responses are obtained along all three axes for the two bracket configurations (aluminium alloy vs hybrid) with packages mounted for the standard random vibration inputs. Sufficient static margin of safety is achieved for high deceleration re-entry flight loads as well. It is concluded that tremendous weight saving on mounting interfaces can be gained without compromising on factor of safety and amplification of vibration responses.

Integrated Navigation, Guidance and Control System and Validation

The navigation, guidance and control (NGC) system of Reusable Launch Vehicle – Technology Demonstrator (RLV-TD) is the most complex system ever flown in ISRO’s launches. It includes a navigation system employing traditional inertial navigation system along with global positioning system aiding and radar altimeter, ceramic servo accelerometer package, angle of attack-based control, different control systems, including secondary injection thrust vector control, reaction control system and aero control surfaces. Closed loop simulations of integrated NGC system of RLV-TD are of utmost importance to ensure mission success by assessment of system performance in nominal and off-nominal flight conditions. Validation is achieved through progressive evaluations in different levels of integrated simulation, namely autonomous simulations, software in loop simulation, on-board processor in loop simulation, hardware in loop simulation and actuator in loop simulation using Iron Bird test facility. This article discusses the overall NGC system, challenges faced during the realization of the simulation test bed, including system design to validate the new on-board elements, commissioning of Iron Bird facility and validation through various phases of simulation. The outcome of these simulations played a crucial role in bringing about improvements and helping in the evolution of the on-board system to the final configuration, ensuring the success of the mission.

Analysis of targeting method for closed-loop guidance of a multi-stage space launch vehicle

This study considers the orbit insertion targeting problem for guidance of a multi-stage space launch vehicle. To design guidance algorithm for a particular mission target, the orbit designation method should be determined first, as the target orbit of closed-loop guidance during the upper stage flight is usually specified at the end of the final stage. We investigate some advantages of this single target approach over the one with intermediate targets in addition to the final target by comparing them in various flight conditions using an explicit guidance algorithm. We notice that in some cases application of the single targeting can be rather tricky and thus propose here an extended version of the single targeting algorithm that can be applied to problems where the target orbit plane has to be changed during the upper stages flight. We also demonstrate that this pseudo-single target method can be applied to these kinds of problems smoothly in contrast to the conventional single target algorithms by conducting numerical simulations for disturbed off-nominal conditions as well as for the nominal condition.

Introduction: Reusable Launch Vehicle Guidance and Control

Introduction: Reusable Launch Vehicle Guidance and Control

Guidance and control of a launch vehicle using a stochastic gradient projection method

A recently developed nonlinear programming stochastic projection of the gradient-type technique is applied to get and test novel procedures for tri-dimensional guidance and control of a satellite solid-fuel launch vehicle. Near optimal open-loop and low-frequency closed-loop solutions are obtained by parametrization of the control history and by considering the satisfaction of boundary constraints unless of specified random errors. Numerical simulations and tests show that the closed-loop solution can cope with the typical model uncertainties and environmental perturbations. A non standard situation of having a fourth stage with a free burning time is tested.

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.

Hybrid approach to near-optimal launch vehicle guidance

Hybrid approach to near-optimal launch vehicle guidance

An expert system for detection, localization and correction of faults in a multi-loop launch vehicle guidance and control system

Abstract A detection, localization and correction of failure in launch vehicle guidance and control system is proposed. The detection and localization of faults is based on monitoring an abrupt change in the error signal in each of the feedback loops, namely, the actuator, digital autopilot, navigation and guidance as well as control loops of inertial sensors such as gyroscopes, inertial platform servos and accelero-meters. Correction schemes are proposed for errors located in each of the loops. A scheme based on reshuffling of control functions to healthy actuators, in the case of a single failure in an actuator loop, is described. Failures in inertia sensors are overcome by software signal reconstruction from healthy sensors. Failures in a guidance loop cause a switchover to an open loop mode that provides vehicle turning rates that are optimal at the last valid loop instant. Retargetting is provided to account for widespread dispersions in vehicle thrust performance.

Optimal explicit guidance of multistage launch vehicle along three-dimensional trajectory

Details of an efficient optimal closed-loop guidance algorithm for a three-dimensional launch are presented with simulation results. Two types of orbital injections, with either true anomaly or argument of perigee being free at injection, are considered. The resulting steering-angle profile under the assumption of uniform gravity lies in a canted plane which transforms a three-dimensional problem into an equivalent two-dimensional one. Effects of thrust are estimated using a series in a recursive way. Encke's method is used to predict the trajectory during powered flight and then to compute the changes due to actual gravity using two gravity-related vectors. Guidance parameters are evaluated using the linear differential correction method. Optimality of the algorithm is tested against a standard ground-based trajectory optimization package. The performance of the algorithm is tested for accuracy, robustness, and efficiency for a sun-synchronous mission involving guidance for a multistage vehicle that requires large pitch and yaw maneuver. To demonstrate applicability of the algorithm to a range of missions, injection into a geostationary transfer orbit is also considered. The performance of the present algorithm is found to be much better than others.

A ring laser gyro based navigator for space launch vehicle guidance

A description is given of the single-package inertial navigation unit (INU) that will replace the three-package gimbaled inertial avionics system in the Centaur upper-stage launch vehicle. The INU was adapted from an existing high-volume, militarized navigation system, based on a ring laser gyro (RLG), that was selected for its performance and adaptability to the space application. Adaptations were made to accommodate: space launch requirement, including qualification in accordance with MIL-STD-1540B; mission accuracies/performance requirements; and vehicle requirements. These adaptations which involved assessment of effects from different operational scenarios and environments, including vibration, thermal, shock, and space radiation, is discussed. Enhancements that have been implemented to comply with INU requirements are examined.< <ETX xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">&gt;</ETX>

Linear-angle solutions to the optimal rocket steering problem

Introduction T optimal guidance of a launch vehicle from a planetary surface into planetary orbit stands out as one of the most successful applications of optimal control theory/calculus of variations. The theoretical foundations of this problem were well developed during the 1950's by a number of investigators, although some isolated and partial work has been done earlier, for example by R. M. Goddard and K.E. Tsiolkovsky in 1903. Lawden's statement of the solution remains as one of the most elegant: