This work presents a modular framework for the design and optimization of pressure-fed engines for liquid-propellant upper stages. Unlike conventional methods, an innovative subalgorithm for the optimization of the inboard profile is implemented. This tool aims to minimize the inert stage mass by identifying the optimal tank configuration—shape, arrangement, and number—that satisfies specific geometric constraints. The framework combines analytical models for the main propulsion and stage subsystems with empirical mass estimation relations to improve the accuracy of mass predictions. A multi-objective optimization strategy, based on a genetic algorithm, minimizes both dry and propellant masses by varying chamber pressure, mixture ratio, and expansion ratio, while respecting mission-derived requirements such as the target ΔV. The analysis of the resulting Pareto fronts highlights clear tradeoffs between propellant mass, structural mass, and engine performance, as well as the influence of geometric parameters like stage radius and length-to-radius ratio on the optimal configurations. Comparison with 10 engines and 6 upper-stage datasets resulted in maximum deviations of 2.7% in dry mass and 3.6% in total mass compared to reference data, suggesting that the tool can reproduce real configurations with very good accuracy.

Well-integrated systems and structures can increase reliability and decrease structural complexity for deployable spacecraft structures. This study investigated the feasibility of a self-deployable space baffle design inspired by the spiral structure of a volute spring. The principles of storing and releasing elastic potential energy within a volute spring were applied to the new baffle design, referred to as the volute baffle. The extension system was integrated with the deployable baffle surfaces to reduce the number of parts, thereby reducing the risk of deployment failure in harsh space environments. A set of derived analytical equations and finite element analysis (FEA) models were used to investigate the mechanical properties of the volute baffle. The linear region of the volute baffle can be accurately modeled by both the FEA and analytical model. The analytical model captured nonlinear behavior slightly better than the FEA solution. A hard stop on extension length was integrated into the spiral structure of the volute baffle, making it more resilient to disturbance and increasing the accuracy of deployed length. Concerns related to static friction, vibration, material choice, and manufacturability were addressed to ensure successful deployments. No deployment failures were observed in the function prototype, which was designed and manufactured with the implemented design features.

An experimental investigation of the hypersonic wake behind the Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) wind tunnel model in a Mach 10 freestream is discussed. Wake visualization was performed using planar laser-induced fluorescence (PLIF) of nitric oxide (NO), which was locally seeded into the wake through the aft body of the model. PLIF measurements are reported at all three of the typical freestream Reynolds numbers of the NASA Langley 31-inch Mach 10 wind tunnel at an α=0 deg angle of attack for two different NO seed gas flow rates. At the lowest Reynolds number, additional data are reported at α=−5 deg and also with the model rolled to φ=90 deg. Unsteady wake boundaries and large-scale unsteady flow structures were clearly seen in the wake at these test conditions. The upper wake boundary was identified using a custom edge-tracking algorithm. Changing the mass flow rate of the NO seed gas by 20× and varying the Reynolds number did not have a significant effect on the average wake boundary, while notable changes in the wake boundary were observed based on the α variation. Data acquired with the model rolled to φ=90 deg provided insight into sting effects.

The space environment is undergoing rapid changes with the emergence of new commercial capabilities, and the development of new technologies to improve space sustainability is imperative. In this context, a comprehensive numerical analysis has been performed to investigate the aerodynamic behavior of a deployable reentry system. In the present investigation, computational fluid dynamics (CFD) and direct simulation Monte Carlo (DSMC) methods are employed to compute the flowfield structure and forces for altitudes ranging from 200 to 95 km. The results indicate that a diffuse high-temperature shock wave is generated upstream of the shield, exhibiting a significant decrease in thickness between 100 and 95 km altitude. At 100 km altitude, good agreement was found between the CFD and DSMC methods, with a difference of 11% in the drag coefficient. This highlights the reliability of CFD within limited altitude ranges and emphasizes the need for DSMC in the upper atmosphere for more realistic results.

Communication blackout is an immense challenge for reentry spacecraft, compromising mission safety. The direct simulation Monte Carlo method with the quantum-kinetic ionization reaction model is used to simulate the ionization flowfield with air–water-vapor injection around RAM C-I (the first flight of the Radio Attenuation Measurement experiments). This study focuses on the three-dimensional flowfield structure and electron number density distribution influenced by water vapor injection, aiming to identify regulation patterns and lay a foundation for further investigation of control mechanisms. Parametric studies are conducted by varying key factors, including the injection pressure ratio, injection velocity, injection temperature, injection angle, injection orifice radius, vehicle’s angle of attack, and the mole fraction of hydrogen ions in the injected water vapor. The elevating effect of the injection is evaluated by comparing the flowfield characteristics across different conditions. Single-factor analysis quantifies the influence of each variable on the electron number density. Through a combined optimization strategy, a reduction of the maximum electron number density at the observation point by one order of magnitude is achieved with a mass flow rate of only 3.38 g/s.

In a pursuit–evasion game, when the evader is unable to escape from the pursuer, the existence of a defender who can defend the evader may change the game outcome. Then, under what conditions (e.g., initial positions, velocities) can the defender guarantee a successful protection? This becomes a significant problem. This study explored this pursuit–evasion–defense game among three coplanar spacecraft, aimed to find the initial orbit conditions that can guarantee the defender–evader (DE) team wins, called winning regions. First, a limit displacement circle method is employed to locate the regions where the defender can arrive earlier than the pursuer, called interception-dominant regions. Based on that, an explicit strategy approach is proposed to transform the solution of the winning region into a cubic or quartic equation regarding the initial state (initial orbital radius ratio, thrust acceleration ratio, and energy constraint). Finally, the approximate analytical expression of the winning region of the DE team was obtained by solving the cubic or quartic equations. Simulation results demonstrate that the computational efficiency of this solution is three orders of magnitude greater than that of numerical methods.

Vision-based navigation is a key technology for spacecraft operating in challenging conditions, such as those encountered at small bodies. To date, most navigation tasks are carried out on the ground, under the supervision of expert operators. However, onboard automation of routine tasks reduces mission costs and at the same time increases the scientific return by avoiding communication delays. In this work, an autonomous vision-based navigation architecture is proposed that exploits two complementary measurements: centroiding and visual odometry. When combined, they provide both absolute and relative navigation information, which can be used to navigate the spacecraft at a variety of ranges. The application scenario is a very close fly-by trajectory on a diverse set of asteroids. The considered targets are characterized by different shapes, sizes, and appearances, highlighting the generality of the proposed navigation pipeline. In-depth testing of the image-processing algorithms is carried out using high-fidelity synthetic images generated in a variety of geometric and illumination conditions. The performance of the navigation architecture is assessed through an extensive Monte Carlo campaign. Results show that the proposed pipeline provides accurate navigation in a wide range of conditions.

One of the most important goals in aerospace engineering applications is the creation of new “flyable” systems. In a flight demonstration using the sounding rocket S-520-34, we show the first successful operation of a bipropellant cylindrical rotating detonation engine using liquid ethanol and liquid nitrous oxide, Detonation Engine System 2 (DES2), in a space environment. From the pressure and temperature histories, the combustion was finished before all propellants were consumed because nitrogen was supplied earlier than the ideal depletion time due to spin stabilization of the sounding rocket. Therefore, the combustion pressure decreased from the nitrogen-supply start time. The short-time Fourier transform result indicated that the deflagration mode, two-wave mode, and single-wave mode occurred in sequence. This was attributed to the locally lower liquid temperatures, wall temperature, and mixture ratio at ignition near the wall, where the rotating detonation wave propagated. A comparison of the filling mass and consumption indicated that the mass flow rate estimated using control surface theory reflects an actual phenomenon. As for the propulsive performance, the experimental characteristic exhaust velocity was almost the same as the ideal value. Moreover, a specific impulse efficiency of more than 90% was achieved throughout the rotating detonation engine operation.

This paper presents the design, development, and flight-test validation of a low-cost, canard-based rocket roll control system, demonstrating an end-to-end process for achieving aerodynamic control capabilities using low-cost, consumer-grade components. The control architecture employs a classical three-loop autopilot framework with time-based gain scheduling, validated using six-degree-of-freedom simulations to ensure stability throughout the rocket’s flight. Designed for cost efficiency and ease of manufacturing, the system incorporates additive manufacturing for various structural components and commercial off-the-shelf servo motors and sensors. Implemented on a Raspberry Pi 4 Model B microcomputer, the system was successfully flight tested on an amateur high-power rocket, demonstrating effective roll control. This work establishes a foundation for a low-cost rocket platform with integrated roll, pitch, yaw, and guidance control, supporting military and commercial research as well as broader operational and developmental applications.

This study presents a detailed computational investigation into the early-stage transient flow behavior of hot-launch missile systems in a concentric vertical tube. Emphasizing the critical milliseconds before thrust buildup and missile liftoff, the work captures the formation and evolution of pressure spikes and shock structures through high-resolution, two-dimensional axisymmetric simulations using a pressure-based solver with the k-ω Shear Stress Transport (SST) turbulence model. The simulations focus on the effects of nozzle inlet pressure and shock interactions on the launch tube wall during the initial seconds of the exhaust gas evolution. Three distinct pressure and temperature peaks are identified, with characteristics resembling Friedlander-type waveforms, driven primarily by bow shock impingement and subsequent wave reflections within the annular gap of the launch tube. Furthermore, the study investigates the influence of nozzle inlet pressure on the pressure and temperature distribution along the launch tube wall. It was observed that while the timing of pressure spike occurrences varied with inlet pressure, the magnitude of the maximum peak pressure remained nearly constant. In contrast, higher inlet pressures resulted in more pronounced temperature spikes, which increased in intensity and propagated further downstream along the launch tube wall. This study offers valuable insights into the design and safety assessment of hot-launch vertical missile systems, aiding the optimization of lightweight launch tubes for high-speed deployment.