Reaction Control System Research Articles

Numerical simulation of influence of jet flow parameters on cone-cylinder-flare configuration

Abstract Reaction control system technology is used to realize attitude control or trajectory change. The hot jet effect should be considered to get more accurate aerodynamic characteristics as the flight vehicle develops. The simulation for hot jet flow needs more calculation resources, so it is important to research the suitable similarity parameters of jet flow. Then, we can use cold jet flow to simulate hot jet flow. In this paper, the influence of jet flow parameters is numerically studied on cone-cylinder-flare configuration. The numerical simulation method for hot jet flow is established, containing finite volume method, multi-component three-dimensional Reynolds Navier-Stokes equation with source terms, multi-block patched grid technique, and fully implicit processing method. The numerical method is verified by using experimental reference results, and the grid independence is assured. The numerical results reveal that the cold jet flowfield under the same mass flow ratio is more similar to the hot jet flowfield than the cold jet flowfield under the total temperature of 293 K. The maximum amplification factor difference between cold and hot jet flow is reduced from 14.1% to 1.2%.

Mass property estimation on TSE(3) via unscented Kalman filter using RCS thrusters

Estimation of spacecraft mass properties in the presence of noise is a nontrivial problem that, when properly addressed, can substantially reduce the crew time devoted to cargo management, control effort expended to manage fuel mass, or both. The estimation of such properties is treated here using a dual unscented Kalman filter (UKF) where the dynamics of the rigid-body spacecraft are formulated on the special Euclidean group SE(3) and its tangent bundle TSE(3) to avoid singularity and nonuniqueness issues found in attitude parameterization sets. The dual UKF on TSE(3) is developed, and a specific methodology is proposed for the estimation of the constant mass properties. An excitation approach is employed in the mass property estimation scheme. Arbitrary sensor suite placements are considered and implemented in the measurement model to capture real-world spacecraft behavior. Furthermore, a reaction control system with duty cycle constraints and noisy thrust values is implemented to account for uncertainty in the excitation inputs. With these features, the methods contained herein are found to be effective at estimating mass properties.

Interaction between lateral jet and hypersonic rarefied flow

Reaction control systems (RCS) are commonly utilized in a variety of air vehicles, and they give rise to complex flow phenomena. For flows dominated by shock waves and expansion, such as lateral jets, a significant gradient is formed in the local region of the flow field, resulting in a high local Knudsen number, and the continuum hypothesis is no longer applicable, even if the freestream belongs to near-continuum flow. As a result, multi-scale simulation approaches should be used to capture the interactions between these multi-scale structures and appropriately determine aerodynamic forces. Our earlier work, which used the conserved discrete unified gas kinetic scheme (CDUGKS) for monatomic gas, studied the flow field features of lateral jets and aerodynamic forces on blunt bodies under various rarefied freestream conditions. In this research, the Rykov model is used to expand the CDUGKS suit for diatomic gas flow while taking into account the excitation of molecular rotational energy. The impact on the flow field and aerodynamic forces from four parameters, namely freestream Mach number, jet Mach number, jet pressure ratio, and nozzle position, is thoroughly investigated. The results show that even if the incoming Mach number is different, the flow field structure and aerodynamic properties don't change much when the jet momentum ratio stays the same. The solid wall stagnation point value and local peak also show some linear relationships. The jet pressure ratio mainly affects the expansion characteristics of the jet, such as the influence range on the solid wall. In addition, with an increase in the distance between the nozzle position and the torque reference point, higher control efficiency can be obtained. The study of these aspects will help us comprehend lateral jet flow and provide reference for the design of an air vehicle's RCS.

Guiding Lunar Landers: Harnessing Neural Networks for Dynamic Flight Control with Adaptive Inertia and Mass Characteristics

The autonomous control of landing procedures can provide the efficiency and precision that are vital for the successful, safe completion of space operations missions. Controlling a lander with this precision is challenging because the propellants, which will be expended during the operations, represent a significant fraction of the lander’s mass. The mass variation of each tank profoundly influences the inertia and mass characteristics as thrust is generated and complicates the precise control of the lander state. This factor is a crucial consideration in our research and methodology. The dynamics model for our lander was developed where the mass, inertia, and center of mass (COM) vary with time. A feed-forward neural network (NN) is incorporated into the dynamics to capture the time-varying inertia tensor and COM. Moreover, the propellant takes time to travel through the feed lines from the storage tanks to the engine; also, the solenoid valves require time to open and close. Therefore, there are time delays between the actuator and the engine response. To take into account these sources of variations, a combined time delay is also included in the control loop to evaluate the effect of delays by fluid and mechanisms on the performance of the controller. The time delay is estimated numerically by a Computational Fluid Dynamics (CFD) model. As part of the lander’s control mechanism, a thrust vector control (TVC) with two rotational gimbals and a reaction control system (RCS) are incorporated into the dynamics. Simple proportional, integral, and derivative (PID) controllers are designed to control the thrust, the gimbal angles of the TVC, and the torque required by the RCS to manipulate the lander’s rotation and altitude. A complex mission with several numerical examples is presented to verify the hover and rotational motion control.

Development of a Ferrofluid-Based Attitude Control Actuator for Verification on the ISS

Ferrofluid-based systems provide an opportunity for increasing the durability and reliability of systems, where mechanical parts are prone to wear and tear. Conventional reaction control systems are based on mechanically mounted rotating disks. Due to inherent friction, they suffer from degradation, which may eventually lead to failure. This problem is further intensified due to the limited possibility for repair and maintenance. Ferrofluid-based systems aim to replace mechanical components by exploiting ferrofluidic suspended motion. Ferrofluids consist of magnetic nanoparticles suspended in a carrier fluid and can be manipulated by external magnetic fields. This paper describes the working principle, design, and integration of a working prototype of a ferrofluid-based attitude control system (ACS), called Ferrowheel. It is based on a stator of a brushless DC motor in combination with a rotor on a ferrofluidic bearing. The prototype will be verified in a microgravity environment on the International Space Station, as part of the Überflieger 2 student competition of the German Aerospace Center. First ground tests deliver positive results and confirm the practicability of such a system.

Systems design results of LunaCube: Dual-satellite lunar navigation system with 6U-CUBESATs

Aiming to establish a long-term presence on the Moon, interest in new and robust communication and navigation capabilities on the lunar surface is greater than ever before. In response, lunar navigation systems (LNS) have been conceptualized and put into action by space agencies. While we expect these services to be fully functional before the late 2020s, many missions are set to begin surface exploration before the initiation of these services. To meet the immediate needs of early missions, our group proposed a small-start technology demonstration mission called the Lunar Navigation CubeSat (LunaCube).The mission aims to demonstrate a LNS transmitter and store-and-forward radio in the vicinity of the Moon and to provide lunar positioning services to lunar surface assets. The project was initiated with funding support from the Japanese Ministry of Education, Culture, Sports, Science, and Technology (MEXT) and completed preliminary design and detailed design in 2021 and 2022 respectively. As a result of the design study, the satellites consist of a 6U-size structure, a 50W-class deployed solar panel, a X-band transponder, three-axis attitude control using a star tracker, sun sensors, gyro, reaction wheels, and reaction control system thruster, as well as orbit control using a water-based resistojet. The total wet mass is less than 14 kg.

Computational study of lateral jet interaction in hypersonic thermochemical non-equilibrium flows using nonlinear coupled constitutive relations

The present study reports the numerical analyses of lateral jet interaction around a Terminal High Altitude Area Defense-type (THAAD-type) model in hypersonic rarefied flows, with the real gas effect incorporated. The computation approach employed is the recently developed thermochemical non-equilibrium nonlinear coupled constitutive relations (NCCR) model. Regarding the simulation conditions, the flight velocity and height are set to 20 Ma and 80 km, respectively. To disclose the flow mechanism of lateral jet interaction, the complex flowfield characteristics and surface pressure distributions are discussed at length. Additionally, the research explores the impact of two key factors, namely, the jet pressure ratio and the jet Mach number, on the control performance of an in-flight vehicle's reaction control system (RCS). The results demonstrate that the complicated flowfield structures in lateral jet interaction are successfully reproduced by the NCCR model. With an increase in either the jet pressure ratio or the jet Mach number, the force and moment amplification factors decrease, while the absolute value of the normal force coefficient increases. Notably, it is found that the rarefied gas effect captured by the NCCR model against the Navier–Stokes–Fourier solution affects the lateral jet interaction flowfield, e.g., weakening the compressibility of the barrel shock and the expansibility of the Prandtl–Meyer expansion fan, as well as strengthening the jet wraparound effect. Importantly, the rarefied gas effect also exerts a prominent influence on the performance of RCS, with the degree of influence diminishing as the jet Mach number or the jet pressure ratio increases.

Six-Degree-of-Freedom Rocket Landing Optimization via Augmented Convex–Concave Decomposition

In this paper an augmented convex–concave decomposition (ACCD) method for treating nonlinear equality constraints in an otherwise convex problem is proposed. This augmentation improves greatly the feasibility of the problem when compared to the original convex–concave decomposition approach. The effectiveness of the ACCD is demonstrated by solving a fuel-optimal six-degree-of-freedom rocket landing problem in atmosphere, subject to multiple nonlinear equality constraints. Compared with known approaches such as sequential convex programming, where conventional linearization (with or without slack variables) is employed to treat those nonlinear equality constraints, it is shown that the proposed ACCD leads to more robust convergence of the solution process and a more interpretable behavior of the sequential convex algorithm. The methodology can also be applied effectively in the case where the determination of discrete controls or decision-making variables needs to be made, such as the on–off use of reaction control system thrusters in the rocket landing problem, without the need for a mixed integer solver. Numerical results are shown for a representative, reusable rocket benchmark problem.

Aftbody Heat Flux Measurements During Mars 2020 Entry

The Mars Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite on the Mars 2020 mission included two total heat flux sensors and one radiometer on the backshell to directly measure the aftbody aerothermal environments during entry into the Martian atmosphere. All three sensors successfully returned aftbody entry heating measurements. Comparisons between the total heat flux sensor measurements and predictions by NASA simulation tools (DPLR/NEQAIR) show excellent agreement and provide confidence in the models. The radiometer measured significant radiative heating, but compared to the model predictions the signal was attenuated by 48% at the end of the entry heat pulse. The loss of signal is attributed to blockage by thermal protection system (TPS) ablation product deposits on the radiometer window. Ground-based testing in the NASA Ames arcjet facilities was conducted to understand the impact of ablation product deposits on the measured radiometer signal. A discussion of the test results, how flight-like the test conditions were, and future work to further characterize the effect of TPS ablation product deposits on the radiometer performance are presented. In addition to measuring the entry heat pulse, all three sensors were sensitive enough to measure solar radiation during cruise, the radiometer measured solar flux during the entry heat pulse, and the leeside total heat flux sensor picked up the descent reaction control system firings.

Uncertainty and Sensitivity Study on Lateral Jet Interaction for Hypersonic Missile

The demanding working environment of hypersonic missiles imposes high requirements on the reliability and stability of the reaction control system. This work aims to investigate the impact of uncertainties in freestream and jet parameters on the flowfield and aerodynamic performance of a missile with a lateral jet. Initially, the influence of angle of attack and Mach number on aerodynamic characteristics is evaluated. The findings indicate that the lateral jet control effect is enhanced as the Mach number increases. Moreover, the interference force/moment does not vary linearly with the angle of attack, and there are notable differences in the flow characteristics when the jet is on the windward and leeward sides. Subsequently, six parameters are selected as input variables. The uncertainties of the output results are analyzed using a stochastic expansion based on the nonintrusive polynomial chaos method. Sensitivity analysis is performed using Sobol indices to estimate the relative contributions of each parameter to the overall uncertainty. Within the given uncertainty interval, the surface pressure of the missile exhibits significant variation, with a maximum uncertainty of approximately 480%. The uncertainty of the recirculation zone range is 50.14%, while the pitch moment amplification factor has an uncertainty of 52.44%. The freestream Mach number, angle of attack, and jet pressure ratio are identified as key factors affecting the aerodynamic performance, surface pressure coefficient, and jet penetration height. Conversely, the contribution of other parameters can be considered negligible.

Numerical simulation of lateral jet interaction with rarefied hypersonic flow over a two-dimensional blunt body

Reaction Control System (RCS) is a direct force control system that successfully adjusts a craft's attitude or orbit using the reaction force created by jet flow. RCS is frequently employed in the management of near-space vehicles due to its properties of fast response time and effective control efficiency. When the near-space vehicle is navigating at high altitude in a low density atmosphere, the Navier–Stokes equation is no longer applicable. The numerical approach utilized in this study is known as the Conserved Discrete Unified Gas Kinetic Scheme, and the governing equation is the Boltzmann equation, which is not constrained by the continuum hypothesis. In velocity space, an unstructured mesh is utilized, which minimizes the amount of discrete velocity points and considerably increases computation efficiency. The numerical results are in good agreement with the direct simulation Monte Carlo code DS2V when modeling large Knudsen number lateral jet flow. The interaction flow field between hypersonic free stream and lateral jet is then simulated at altitudes of 60–90 km using argon as the working gas and a two-dimensional blunt cone with lateral jet as the study object. Under a fixed jet pressure ratio, preliminary research was conducted on the variation of the lateral jet interference flow field characteristics with the freestream Knudsen number and angle of attack. The differences in surface pressure and heat flux caused by jet opening and shutting are compared. Under rarefied atmospheric conditions, the variation of the force/moment amplification coefficient is given. The numerical results show that when the angle of attack is 0°, the separation area in front of the nozzle and a pair of opposite vortices, which are common in the jet interference flow field, gradually disappear with increasing altitude, but the separation vortex reappears when the angle of attack of the freestream is increased. The high-pressure region generated upstream of the nozzle is the primary cause of the extra force/moment. The density of the main flow decreases as altitude increases, various shock wave patterns of the interference flow field gradually dissipate and the force/moment amplification factor changes considerably. The rarefied gas effect has a significant effect on the lateral jet interference flow.

Fixed-time attitude control for aircraft with strongly constrained actuators

This paper investigates a novel fixed-time attitude control method for trans-medium aircraft to ensure proper contact between the fuselage and the new medium before transmitting from different environmental mediums. How to achieve rapid corrective attitude adjustment from improper contact by utilizing the Reaction Control System with multiple discrete thrust gears that possess discrete thrusts with limited output has become a great challenge. To solve this problem, a novel fixed-time attitude controller is proposed. Firstly, a non-singular segmented terminal sliding mode surface is constructed to achieve a fixed convergence time and avoid the singularity problem. Subsequently, a command regulator is designed to cope with the derivation introduced by the strongly constrained actuators, which consists of the approximate thrust model and an auxiliary system based on a double power function. Furthermore, the stability analysis indicates the fixed time convergence stability of the closed-loop system and provides a clear calculation of the settling time. Finally, numerical comparison results and Monte Carlo simulations verify that the attitude can converge in a predetermined time from different initial states.

Development of high burn rate propellant and testing in miniature rocket motor for control applications

For aerospace vehicles, notably missiles or satellites for attitude and divert control, impulsive thrusters are most desirable if they are quantized as exact instantaneous pulses available over and after extended periods of inactive storage. These thrusters operate for a few milliseconds (ms), typically between 10 ms to 100 ms duration. Compared to its liquid or gaseous counterpart, an impulsive thruster powered by solid propellants is a simple, and reliable solution. Therefore, they are suitable as reaction control systems and divert thrusters for satellite and unmanned aerial vehicle applications.A typical solid propellant would need an extremely thin web thickness for port burning arrangement due to its brief burn duration, which makes propellant manufacturing and its structural anchoring in motor challenging. A high burn rate and high-density propellant based on ammonium perchlorate, aluminum, and Teflon is processed to enable end-burning configuration thus solving the problems, simplifying propellant manufacturing, structural design and integration. It also permits stacking propellants for multi-pulse applications. The high burn rate (40 mm/s at 120 bar pressure) propellant is formulated and tested for ballistic performance evaluation in proof motor configuration, and evaluation of its advantages over conventional propellant. A case of identical design requirements with both double base propellant and high burn rate pressed propellant is compared, tested and evaluated to compare the relative performance merits. The volumetric load is increased from 47% to 82% due to the development and adoption of high burn rate propellant. The overall weight of the thruster is also reduced by 20%.

Method of 2-D Range-Doppler Imaging for Plasma Wake Based on Range Walk Correction

When a hypersonic vehicle flies in the atmosphere, plasma sheath and trails are configured around the vehicle due to aerodynamic heating and ablation, which distort the radar echo signal and then influence the target range information and Doppler frequency extracted from the echoes. In order to examine the impact of plasma laminar wake on radar detection, this article is primarily aimed to establish a laminar wake echo model based on a linear frequency-modulated continuous wave (LFMCW) radar, which takes into account the reaction control system of laminar wake by employing the Born approximation solution and the characteristics of the inhomogeneous and nonunique velocity of the wake medium. An echo signal processing method is also developed, in which the stretch processing technique is combined with the keystone transform of the target range walk correction. Such a treatment not only reduces the sampling rate and bandwidth, but also eliminates the cross-range cells phenomenon caused by hypersonic; finally, the range-Doppler (RD) image of the wake is achieved by 2-D fast Fourier transform (2-D FFT) processing, and radar range and velocity measurement results in the presence of plasma are obtained.

Propulsion Systems, Propellants, Green Propulsion Subsystems and their Applications: A Review

A wide range of propellants, and propulsion systems in space exploration by aircrafts or space vehicles was studied, developed, investigated, and commercialized. Liquid, solid, or hybrid propellants have been used for rocket’s launches. In this review, a consistent definition of space propulsion systems, including solid, liquid and hybrid has been given with up-to-date state of developments. A comparison of their performances was made by theoretical and experimental specific impulses. On the other hand, ammonium perchlorate and hydrazine were used as propellants for rocket’s launches and for satellite’s maneuverings; respectively. However, their high toxicity and their storage problem pushed researchers and scientists to investigate and develop other eco-friendly, propellant systems, so called “green propellants”, for launch or reaction control systems of satellites.

Path-Following Control for Thrust-Vectored Hypersonic Aircraft

Thrust vector control (TVC) might be used to control aircraft at large altitudes and in post-stall conditions when aerodynamic control surfaces are ineffective. This study demonstrated that the implementation of the TVC on high-speed aircraft is a reasonable solution and might be an alternative when compared to the complicated reaction control system or large aerodynamic control surfaces. The numerical flight dynamics model of the X-15 experimental aircraft was developed and implemented in MATLAB/Simulink and then used to investigate the proposed solution. The obtained results indicate that the aircraft, equipped with full 3D thrust vectoring and two independent horizontal stabilizers to control the roll angle, was able to achieve flight along the path that was defined by a set of waypoints. This paper also highlights the potential benefits and challenges of using TVC as a control method for aircraft. The results of this study contribute to the growing body of research on aircraft control and simulation. Future work can explore the use of TVC for other aircraft with unique configurations and low maneuverability features.

Predictive Control for Large Angle Attitude Tracking Maneuver of Spacecraft with Reaction Control System and Reaction Wheel

Missions involving rapid and large angle attitude maneuvers have been conceived in astronomical and earth observation satellites in recent years. As an actuator capable of generating a large torque is also going to be required, it will also be necessary to consider characteristics of an actuator in designing a control system. Actuators capable of generating a large torque include reaction control system (RCS). RCS gives an on/off input as it uses the reaction force from fuel injection by thrusters, it can generate a large moment. In addition, the control system of current application satellites normally uses both RCS and reaction wheel (RW) conventionally used for attitude control. This paper considers large angle attitude maneuver of spacecraft by a combination of RCS and RW. To this end, characteristics of RCS and RW are defined, and a discrete-time model (Euler approximation model) for the control system design is derived. Next, we design a discrete-time nonlinear tracking controller so that the closed-loop system consisting of the Euler approximation model becomes input-to-state stable (ISS) using the concept of backstepping approach. Then, we propose a method to appropriately change the threshold of RCS injection corresponding to the attitude of spacecraft by model predictive control. Finally, the effectiveness of proposed control method is verified by numerical simulations.

Autonomous Wheel Off-Loading Strategies for Deep-Space CubeSats

Deep-space CubeSats missions require careful trade-offs on design drivers such as mass, volume, and cost, while ensuring autonomous operations. This work elaborates the possibility of off-loading the reaction wheels without the need of carrying a bulky and expensive reaction control system or the field-dependent magnetotorquers. The momentum accumulated along two body axes can be removed by either offsetting the main thruster with a gimbal mechanism or by tilting differentially the solar wings. The dumping on the third axis can be still accomplished by imposing a specific attitude trajectory with the motion of either the gimbal or the arrays drive mechanism. The M-Argo CubeSat is selected as case study to test the techniques along its deep-space trajectory. The strategies decision-making is autonomously carried out by a state machine. The off-loading during the cruising arcs employs the gimballed thruster and takes typically 3 h, granting a mass savings of more than 99% with respect to the usage of a reaction control system. The trajectory is shown to have negligible differences with respect to the nominal one, since the thrust is corrected accordingly. During the coasting arcs, the solar arrays are tilted and several hours are required, depending on the Sun direction and intensity, but the propellant is completely saved. Sensitivity analyses are also carried out on the initial angular momentum components and the center of mass displacement to check the robustness of the algorithms.

Performance Evaluation for Launcher Testing Vehicle

The paper’s purpose is to present a calculus model for a testing vehicle that can be used to validate guidance, navigation and control systems for reusable launchers in all flight phases. The technical solution is based on a throttleable engine with thrust vectoring control and a reaction control system (RCS) used for roll. For calculus, we will develop a nonlinear model with six degrees of freedom, based on quaternion, extended with nonlinear equations that use pulse modulation in order to control roll. In order to synthesize the controller, we also develop a linear model similar to the launcher model. The paper analyzes two basic scenarios, first with the ascending and the descending flight phases and the second having a horizontal flight interleaved between ascending and descending flight phases, both scenarios being specific for reusable launchers. Based on these scenarios, the paper evaluates some performances of the proposed vehicle, namely flight envelope and guidance accuracy.

Self‐compensation of torque reaction in an induction contra‐rotating propulsion motor

Torque reaction is one of the most critical issues with current aerial and undersea vehicles. Contra-Rotating Propulsion (CRP) systems are the most popular technique for compensating the torque reaction of floating vehicles. However, the unbalanced load conditions are inescapable in CRP systems, which eliminate the net zero torque reaction and impulse unwanted spinning on vehicle. A new single-phase dual rotor axial flux induction motor has been investigated in this paper that has the ability of self-compensate reaction torque against unbalanced loads. The proposed motor can eliminate the requirement for complicated torque reaction control systems. Finite element simulation shows the expected performance of proposed motor under unbalanced conditions. The simulation results have been verified by experimental testing of a motor prototype.