Thrust Vector Control for Nuclear Thermal Rockets

The controlled atmospheric re-entry associated with the precision soft-landing of Reusable Launch Vehicles (RLVs) on Earth is very challenging as it depends on multiple parameters [1]. Over the last decade, the cost-effectiveness of such a technology has been finally demonstrated with the successful recoveries of SpaceX’s Falcon 9 first-stage rocket first [2], then followed by other companies such as the Rocket Lab’s Electron micro-launcher [3]. This breakthrough has been made possible by the development of advanced and robust computational methods able to generate in real time the flight conditions and to command the optimal vehicle's deflections accordingly to achieve a safe pinpoint landing. Indeed, during an Earth atmospheric re-entry, the vehicle is subjected to fast system dynamics changes partly induced by external loads associated with the terrestrial environment (e.g., lift, drag, wind and gusts), but also by the actuation commands to answer the landing constraints satisfaction and the vehicle integrity preservation. All those involve uncertainties and nonlinearities, which lead to vehicle’s instability and therefore give reason why for a highly performant Guidance, Navigation and Control system implementation [4]. More particularly, one of the critical aspects is the design of a robust control strategy capable of counteracting the previously defined disturbances and uncertainties while satisfying the strict accuracy requirements associated with the pinpoint landing [5]. As demonstrated by the current state-of-the-art on control design for launchers [6-7], the classical linear control theory represents a rich heritage with a lot of applications. This choice was motivated by its relative easiness of implementation and the possibility to use gain-scheduling techniques to adapt to nonlinear systems. Nevertheless, these techniques are well-adapted to the control system design of single-input single-output systems, such as for example a reusable rocket using a Thrust Vector Control (TVC) system as the unique actuator. The implementation of multiple-input multiple-output control systems becomes then complex since every channel is addressed in a single-loop fashion. This capability is however required for the future generation of reusable rocket, using also aerodynamic steering based on fins in addition to the TVC system for enhancing control authorities. Moreover, model uncertainties are not accurately considered in the design process, developed only with nominal conditions and stability margin requirements. For all these reasons, it results in an extensive (both in terms of time and cost) Verification and Validation campaign with many iterations and Monte-Carlo analyses to assess the performance and robustness of the control system. To overcome these drawbacks, the H-infinity family of methods, introduced a few years ago [8], provides with a powerful solution for robust control design. It relies on defining the control requirements in the frequency domain in terms of weighting functions and minimising the maximum gain of the resulting weighted system from the exogenous inputs to the outputs to be controlled. Moreover, the control-plant interaction is modelled through a Linear Fractional Transformation (LFT) representing the feedback action. Finally, the structured H-infinity method [9] allows to directly impose a specific control structure – like a Proportional-Integral-Derivative (PID), enabling the re-use of gain-scheduling techniques – and to consider parametric uncertainties for enhanced robustness. This paper studies the synthesis of a robust control system via structured H-infinity for the RLV atmospheric re-entry problem. First, the nonlinear 6-Degree-of-Freedom (6-DoF) RLV re-entry dynamics are simplified into a linear model and then linearised along a reference trajectory to get the nominal LFT of the system, then augmented with parametric uncertainties. The model covers the atmospheric re-entry and vertical landing of a first-stage rocket equipped with a TVC system and steerable planar fins. The controllers are built at different points of the re-entry trajectory, using the structured H-infinity framework through PID-like structures. Weighting functions considering the control objectives and requirements of a realistic RLV re-entry scenario are implemented. A robust stability analysis of the obtained system is performed via classical stability margins and structured singular value. Finally, the controllers are gain-scheduled and validated via Monte-Carlo analyses, using a nonlinear 6-DoF RLV re-entry dynamics simulator equipped with successive convex optimisation guidance. The performance of the resulting guidance and control architecture is compared with the baseline developed in previous works [10], using a TVC system only and classical linear feedback control through gain-scheduled PID controllers. This study lies within the ASCenSIon (Advancing Space Access Capabilities - Reusability and Multiple Satellite Injection) project, an innovative training network funded within H2020.

In this research, the effects of some liquid side injection from nozzle wall into exhaust gas of combustion chamber are studied. The side injection against main flow is as elliptical solid thing that change the symmetric of flow field on nozzle wall and causes some different pressure distribution on wall, and finally causes thrust vector deviation. Flows interaction causes some physical phenomena as bow shock wave in front of injection region. This paper explain the effects of this wave and variation velocity & pressure distribution at different cross sections of flow field and comparison results of air and other liquid fluid in thrust vector control system. The results are compared with experimental data and have well agreement with them. The results show that Freon is one of best injection liquid for this type of thrust vector control. Performance of Injection is optimum in relative position 35 to 40% nozzle divergence length.

Thrust vector control (TVC) methods were developed as an alternative to aerodynamic control surfaces for missiles, which suffered from poor pitch-over characteristics in launching stage. TVC system, can allow the missile to turn quickly and stably in any direction by deflecting the nozzle exhaust gas irrespective of the missile speed. A jet vane is one of such devices, used for thrust vectoring. In conventional approaches, jet vane is connected control surface through a mechanical interface such as wire and link. However, these approaches lead to angular position dependent transmission ratio. In addition, these techniques limit designer to obtain the required hybrid control ratio. In this study, a novel compact electromechanical actuator is proposed to overcome these problems. Constant or variable transmission ratio can be obtained using such actuator. Proposed solution is not only a simple and compact structure, but also being able to provide sufficient control torque for a missile's quick turn.

Mars mission capabilities enabled by nuclear thermal propulsion

This paper outlines the development and implementation of a Thrust Vector Control (TVC) system specifically designed for solid propellant rocket engines, aiming to enhance the portability of hybrid and liquid propellant rocket systems. The proposed system ensures trajectory stability and rapid response to flight emergencies by effectively counteracting external perturbations such as wind. Key components include gyroscopic (GYRO) sensors, a high-performance microcontroller, and servo motors, collectively referred to as thrust vectoring components, which dynamically adjust thrust direction opposite to the rocket’s trajectory to maintain stability. Drawing from an extensive literature review of existing TVC system designs, with a focus on thrust characteristics and engine combustion timing, the design integrates geometric properties, kinematics, forces, energy requirements, safety, cost, and control methods, aligning with both mechanical and avionic design principles. The anticipated outcome is an advancement in rocket propulsion technology, characterized by improved angular speed control, trajectory linearity, and rapid response capabilities.

Actuators are used to provide Thrust Vector Control (TVC) for lifting the launch vehicle and its payload to reach its space orbits. The TVC comprises actuation and power components for controlling the engines direction and thrust to steer the vehicle. Here the actuator considered is a linear Electromechanical Actuator (EMA) system. The design is based on the Model Reference Adaptive Controller (MRAC) technique. It calculates an error value as the difference between actual system and desired model and attempts to minimize the error by adjusting the controller parameters. Adaptive control laws are established with the gradient method. The experimental result shows that the MRAC has better tracking performance than that with the traditional compensated system when both controllers are subjected to the same parameter variations. The simulation is done using MATLAB/SIMULINK software.

At the current stage of space exploration development, missions to the Moon and geostationary orbit are relevant. Enhancing of power supply level of such missions requires innovative technology development. Nuclear rocket engines and engines with use of solar energy are considered among such technologies. In a number of projects, such engines use intermediate heat energy accumulators, which complicates and increases the cost of construction. This embodiment uses a circuit that enables solar engine to operate with any orientation of the thrust direction towards the sun. The dependence of specific thrust of such engine on maximum heating temperature of working fluid (hydrogen) is shown. Such engines provide the ability to return to Earth at a speed of 8 km/s instead of 11 km/s, which increases flight safety and makes propulsion system reuse efficient.

A study was undertaken to determine if movable probes or struts positioned in the nozzle can be used to provide Thrust Vector Control of the Space Shuttle Solid Rocket Booster. The study employed CFD to determine estimates of the shock standoff distance from the probe. An empirical correlation was used to construct the shock shape and the pressure distribution generated by the probe. The TVC performance for a single and multiple number of probes was then used to determine requirements for a maximum thrust angle offset of 7.5 degrees. Consideration was given to what materials would be suitable for the probe and if active cooling is required. Based on the performance analysis and thermal requirements, a Probe Thrust Vector Control (PTVC) system was sized. Indications are that a PTVC system weight is in the 1500 1bm weight range, compared to the existing weight of 7500 1bm for the SRB nozzle gimble system.

Thrust vector control technology is widely used in modern advanced aircrafts, space shuttles and unmanned aerial vehicles, etc. An accurate thrust vector mathematical model is of great significance to enhance the thrust vector control effect. In this paper, the P200 micro-turbine engine with an axisymmetric thrust vector nozzle placed at the rear is taken as an example, A hybrid thrust vectoring dynamics modeling method consisting of engine speed model, engine thrust model, actuator dynamics model, airflow angle model and thrust loss coefficients is proposed. The engine speed model is of Linear Parameter Varying (LPV) model with each segment is characterized by a First Oder Plus Dead Time (FOPDT) model. The engine thrust model is characterized as a cubic polynomial function of engine speed. The actuator’s dynamic model is a FOPDT model. The nozzle airflow deflection angle is proportional to the airflow deflection angle. the thrust loss coefficient is approximated by a cosine relationship with the airflow deflection angle. The experimental results show that the established model can well characterize the dynamics of the thrust vector system.

A study was initiated to investigate propulsion stage and mission architecture options potentially enabled by fission energy. One initial concept is a versatile Nuclear Thermal Propulsion (NTP) system with a maximum specific impulse of 900 s and a maximum thrust (per engine) of 15 klbf. The system assumes a monopropellant stage (hydrogen), and is designed to also provide 300 lbf of thrust (potentially split between multiple thrusters) at an Isp > 500 s. Boost pumps are used to assist with engine decay heat removal and low thrust engine burns, and to compensate for partial tank depressurization during full thrust engine burns. Potential stage assembly orbits that take full advantage of launch vehicle payload mass and volume capabilities are being assessed. The potential for using NTP engines to also generate a small to moderate amount of electrical power is also being evaluated.

The Nuclear Thermal Rocket is an enabling technology for human exploration missions. The “bimodal” NTR (BNTR) provides a novel approach to meeting both propulsion and power requirements of future manned and robotic missions. The purpose of this study was to evaluate tie‐tube cooling configurations, NTR performance, Brayton cycle performance, and LOX‐Augmented NTR (LANTR) feasibility to arrive at a point of departure BNTR configuration for subsequent system definition.

According to the thrust vector control requirements of fast turn of the missile vertical launching, this design carries out the force analysis of the jet vane which is in the high temperature and high speed jet stream, and calculates the normal force which is generated by the vane, thus calculates the area of the jet vane. At the same time, according to the aerodynamic theory and engineering experience, this design analyzes and calculates the aerodynamic shape of the jet vane. As a result, two types of jet vane are designed. On this basis, through the use of a fluid analysis software FLUENT, both of the two jet vane are simulated and get a jet vane which is more suitable for the practical applications.

The thrust vector control problem for an upper‐stage rocket with propellant slosh dynamics is considered. The control inputs are defined by the gimbal deflection angle of a main engine and a pitching moment about the center of mass of the spacecraft. The rocket acceleration due to the main engine thrust is assumed to be large enough so that surface tension forces do not significantly affect the propellant motion during main engine burns. A multi‐mass‐spring model of the sloshing fuel is introduced to represent the prominent sloshing modes. A nonlinear feedback controller is designed to control the translational velocity vector and the attitude of the spacecraft, while suppressing the sloshing modes. The effectiveness of the controller is illustrated through a simulation example.

Future space missions, such as return from Mars, will require entry speeds of the order of . These harsh conditions are dominated by radiative heat flux and have large modeling uncertainties related to radiative coupling, such as precursor radiation and ablation radiation coupling. These conditions have also never been tested in a flight scenario and have limited ground testing data available. This study investigates the feasibility of generating these conditions in the X2 expansion tube at the University of Queensland so that these flows can be tested over scaled test models for the study of shock relaxation, heat flux, shock standoff, or ablation. This paper presents a theoretical performance map of the X2 facility for the generation of these conditions before it presents some of the unique issues that may be encountered when generating and investigating them. Some preliminary experimental data are then presented, showing that these 15 km/s equivalent flight speed test conditions can be generated in X2 within the current performance limits of the facility.

<div class="htmlview paragraph">A major effort at the NASA Lewis Research Center in recent years has been to develop electromechanical actuators (EMA's) to replace the hydraulic systems used for thrust vector control (TVC) on launch vehicles. This is an attempt to overcome the inherent inefficiencies and costs associated with the existing hydraulic structures. General Dynamics Space Systems Division, under contract to NASA Lewis, is developing 18.6 kW (25 hp), 29.8 kW (40 hp), and 52.2 kW (70 hp) peak EMA systems to meet the power demands for TVC on a family of vehicles developed for the National Launch System. These systems utilize a pulse population modulated converter and field-oriented control scheme to obtain independent control of both the voltage and frequency. These techniques allow an induction motor to be operated at its maximum torque at all times. At NASA Lewis, we are building on this technology to develop our own in-house system capable of meeting the peak power requirements for an expendable launch vehicle (ELV) such as the Atlas. Our EMA will be capable of delivering 22.4 kW (30 hp) peak power with a nominal of 6.0 kW (8 hp). This system differs from the previous ones in two areas: 1) the use of advanced control methods, and 2) the incorporation of built-in-test. The advanced controls are essential for minimizing the controller size, while the built-in-test is necessary to enhance the system reliability and vehicle health monitoring.</div> <div class="htmlview paragraph">The ultimate goal of this program is to demonstrate an EMA which will be capable of self-test and easy integration into other projects. This paper will describe the effort underway at NASA Lewis to develop an EMA for an Atlas class ELV. An explanation will be given for each major technology block, and the status of the overall program will be reported.</div>