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In this paper, a novel integrated guidance and control algorithm based on adaptive high-order sliding mode is proposed for reusable launch vehicle subject to unknown disturbances and actuator faults. We propose a time-varying barrier function-based adaptive control law to offset the effects of uncertainties. The remarkable feature of the developed algorithm is its ability to track the reference commands in finite time despite unknown disturbances and actuator faults, without designing the guidance law and attitude controller separately. Finally, the effectiveness of the proposed algorithm is confirmed by the simulation results.

Recently, in the space market, there has been a rapid reduction in the launch price. The major reason for this is that a few commercial companies began to enter the space market about ten years ago, which has changed the space market from monopolization to competition and accelerated the adoption of commercial efficiency in technology and management. Also, SpaceX made a breakthrough by successfully recovering the first stage of its launch vehicle in 2016, thereby opening the door to reusable launch vehicles. They have declared their intention to significantly reduce satellite launch costs in the future by utilizing the potential of reusable launch vehicles. In this study, we calculate the total launch cost required to place a single satellite into Low Earth Orbit (LEO) and compare the launch costs for three different payload capacity scenarios, investigating how payload capacity affects the cost-effectiveness of reusable launch vehicles. The launch cost is broken down into development costs, production costs, reuse costs, operational costs, fixed costs, and insurance costs, with cost estimation equations utilized based on cost calculation models such as TRANSCOST.

This paper presents the network-based modeling, validation and analysis of the nonlinear liquid spring damper model under vertical landing conditions of reusable launch vehicle. The impedance function of damper model is derived first. Then, its mechanical and hydraulic networks are newly established based on the hydro-mechanical analogy and network-based analysis. By comparing the networks between the corresponding symmetric and asymmetric structures, the meaning of each branch in the network is elucidated. After that, the validity of the network-based model for the liquid spring damper is confirmed by comparison against the experimentally verified nonlinear model in both frequency and time domain. The force and energy absorption characteristics of the damper model are further decomposed, and, specifically, the influence of the orifice area and orifice length on the attenuation performance is studied. The results show that the network-based model provides predictions consistent with those generated by the nonlinear model. The main discrepancy is attributed to the inaccuracy caused by the equivalent fluid bulk modulus. The network-based analysis indicates that the orifice area mainly influences the damping force in the network, which further affects the loads and efficiency of the damper. The orifice length mainly influences the inertia force in the network, which should be limited to a small value. The proposed novel interpretation of the damper models and responses under impact conditions constitutes a framework suitable for systematic design of typically highly nonlinear landing systems in reusable launch vehicles.

With the cost-saving benefits of reusable launch vehicles (RLVs), South Korea is pursuing the application of reusability technologies to KSLV-III. While SpaceX currently reuses only the first stage of Falcon 9, the Starship program aims for full-stage recovery, motivating further examination of second-stage reuse. This study extends the scope of the analysis to medium-class launch vehicles and evaluates the feasibility of second-stage reuse for two vehicle scales. The performance losses associated with three recovery methods—vertical landing, parachute, and fly-back—are quantitatively assessed using conceptual-level recovery system design and simplified mass modeling. For KSLV-III, a conceptual expendable medium-class launch vehicle capable of delivering a 10-ton payload to a 200 km low Earth orbit (LEO) was designed using an algebraic modeling approach. Based on this reference design, the recovery methods were evaluated for both medium-class and super-heavy-class vehicles. Results of the present order-of-magnitude conceptual trade study show that, for medium-class vehicles, the parachute provides the highest performance, followed by fly-back, while vertical landing yields the lowest. For super-heavy vehicles, the parachute remains the most effective, but vertical landing becomes the second-best option, with fly-back exhibiting the lowest performance. As the vehicle scale increases, parachute effectiveness declines, fly-back performance improves, and vertical landing shows the greatest performance gains. However, parachute becomes impractical for super-heavy vehicles due to structural limitations, making vertical landing the most viable option. In contrast, medium-class vehicles do not necessarily require vertical landing, and the optimal recovery strategy should be chosen based on vehicle structural characteristics and mission objectives. This study provides insights that support the selection of efficient recovery strategies during the early design phase of RLVs.

For the diverse payloads of Reusable Launch Vehicles and the inevitable problem of change in the center of mass, this paper proposes an incremental nonlinear dynamic inversion considering centroid variation control. Regarding the trans-atmosphere flight environment, the six-degree-of-freedom dynamics model considering centroid shift, Earth rotation, and the Clairaut Ellipsoid Model is established to improve model accuracy. An incremental nonlinear dynamic inversion considering a centroid variation controller with excellent dynamic performance and adjustment under the centroid variation is designed for the model, which fully meets the safety requirements of RLV reentry. An extended state observer considering centroid variation is proposed to solve the problem with difficult direct measurement of angular acceleration, which incorporates the influence of centroid variation into the known part to improve estimation accuracy and speed. Finally, the simulation results are provided to verify the robustness of the change of centroid position and good control quality with the proposed controller.

After the successful operationalization of Polar Satellite Launch Vehicle and Geo Synchronous Launch Vehicle, the Indian Space Research Organisation is in the process of developing Reusable Launch Vehicle technologies to achieve low-cost access to space. Towards this programme, a winged body configuration was conceived, which can fly at subsonic, supersonic and hypersonic Mach number regime, re-enter into the earth's atmosphere and simulate the landing manoeuvre. The aerodynamic design, analysis and wind-tunnel testing, aerothermal and structural design, analysis and testing were carried out. Suitable solid motor with slow burn rate propellant was developed. Mission design, guidance and control schemes were implemented. In order to meet the above objectives, certain technologies and infrastructure were developed. The entire subsystems were integrated and a large number of flight measurements were made in the maiden successful flight of Reusable Launch Vehicle Technology Demonstrator in May 2016. The flight measurements and flight performance indicated that the design philosophy, testing schemes and approaches followed are in order, thus providing confidence to proceed to the next logical step in the development of Reusable Launch Vehicle Technologies.

In this work, a vertical landing mechanism of a reusable launch vehicle (RLV) is investigated using a flexible–rigid coupled dynamics model. The presented model takes into account the four-legged landing mechanism and the main body cabin. Flexibilities of the main components in the vertical landing mechanism are considered. The hydro-pneumatic spring force and thrust aftereffect caused by the sequential deactivation of the engine are introduced separately. Several simulation cases are selected to analyze the loads acting on the landing mechanism and the dynamics behavior of the whole RLV system. Simulation results show that considering flexibility in the landing mechanism is critical for dynamics analysis under various initial conditions. The adopted RLV design is capable of achieving stable landings under specified initial velocity and attitude conditions, demonstrating its feasibility for engineering applications. Moreover, the hydro-pneumatic spring plays a crucial role in absorbing the impact of the initial landing leg, ensuring a smoother landing experience and minimizing potential damage to the vehicle.

During the reentry terminal flight of lifting-body Reusable Launch Vehicles (RLVs) propelled by liquid fuel, the sloshing of liquid propellent presents new features that, if neglected, could lead to adverse flight oscillations or even worse. This paper focuses on liquid sloshing coupled flight dynamics, sloshing effect prediction, and the suppression of adverse flight oscillations. First, a transfer function model for unsteady aerodynamics is improved and applied to describe the sloshing force effect, being included in the rigid–liquid control coupled flight dynamics model. The frequency domain analysis results show that liquid sloshing tends to degrade the closed-loop stability margin of the vehicle and even induce less damped oscillations, which can be predicted through the frequency characteristics with the sloshing force effect included. Furthermore, three suppression control measures to mitigate adverse oscillation are addressed, which include enhancing the trajectory-tracking loop damping, separating the frequencies of the rigid body motion and the liquid sloshing, and especially introducing a compensation loop to counteract the sloshing effect. Simulations demonstrate that all the provided approaches help mitigate the sloshing effect, while the compensation control with sloshing frequency characteristics included works best.

Purpose This paper aims to present a novel adaptive sliding mode control (ASMC) method based on the predefined performance barrier function for reusable launch vehicle under attitude constraints and mismatched disturbances. Design/methodology/approach A novel ASMC based on barrier function is adopted to deal with matched and mismatched disturbances. The upper bounds of the disturbances are not required to be known in advance. Meanwhile, a predefined performance function (PPF) with prescribed convergence time is used to adjust the boundary of the barrier function. The transient performance, including the overshoot, convergence rate and settling time, as well as the steady-state performance of the attitude tracking error are retained in the predetermined region under the barrier function and PPF. The stability of the proposed control method is analyzed via Lyapunov method. Findings In contrast to conventional adaptive back-stepping methods, the proposed method is comparatively simple and effective which does not need to disassemble the control system into multiple first-order systems. The proposed barrier function based on PPF can adjust not only the switching gain in an adaptive way but also the convergence time and steady-state error. And the efficiency of the proposed method is illustrated by conducting numerical simulations. Originality/value A novel barrier function based ASMC method is proposed to fit in the amplitude of the mismatched and matched disturbances. The transient and steady-state performance of attitude tracking error can be selected as prior control parameters.

The vertical-takeoff–horizontal-landing (VTHL) two-stage-to-orbit (TSTO) system is a kind of novel launch vehicle in which a reusable first stage can take off vertically like a rocket and land horizontally like an airplane. The advantage of the VTHL TSTO vehicle is that the launch costs can be reduced significantly due to its reusable first stage. This paper presents an application of multidisciplinary analysis optimization on preliminary sizing in conceptual design of the VTHL TSTO vehicle. The VTHL TSTO concept is evaluated by multidisciplinary analysis, including geometry, propulsion, aerodynamics, mass, trajectory, and static stability. The preliminary sizing of the VTHL TSTO vehicle is formulated as a multidisciplinary optimization problem. The focus of this paper is to investigate the impacts of the first-stage reusability and propellant selection on the staging altitude and velocity, size, and mass of the VTHL TSTO vehicles. The observations from the results show that the velocity and altitude of the optimal staging point are determined mainly by the reusability of the first stage, which in turn affects the size and mass of the upper stage and the first stage. The first stage powered by hydrocarbon fuel has a lower dry mass compared with that powered by liquid hydrogen.