Future projections indicate that continued population growth will lead to further expansion and densification of urban environments, thereby increasing transportation demands and associated challenges. In this context, Urban Air Mobility (UAM) has emerged as a promising solution, enabling new intra- and inter-urban transportation services through the use of Vertical Take-Off and Landing (VTOL) aircraft, more precisely configurations such as lift and cruise tiltrotors which combine the hovering capability of conventional helicopters with the cruise speed and range of fixed-wing aircraft by means of tilting propulsion mechanisms. Optimizing the aircraft design process is essential to reduce overall development time and cost. During the conceptual design phase, propeller design methodologies commonly reported in the literature rely on vortex-based approaches or actuator disk theory to estimate the main propeller characteristics. However, the accuracy of these methods strongly depends on the inflow angle and operating conditions, with discrepancies increasing as the inflow angle and advance ratio grow. This paper introduces an analytical model to predict propeller thrust at a 90° inflow angle (pure lateral flow), based on a correction of the thrust under perpendicular flow conditions and the propeller geometry evaluated at 75% span. The approach relies on local velocity and angle of attack estimations derived from classical Blade Element Momentum Theory (BEMT). The propeller lift coefficient is obtained by representing the blades as thin airfoils, with an additional correction to account for stall effects at high angles of attack, while the drag coefficient is calculated based on the known thrust values. The induced velocity, required for local flow calculations, is estimated from known thrust values and discretizing the propeller disk according to the number of blades. This capability is particularly relevant for modeling lift and cruise tilt rotor configurations cruise phase during early design stages while maintaining minimal computational cost. The proposed model is validated against wind tunnel measurements for several propellers tested at different global pitch angles, demonstrating the applicability of the developed formulation for blades with twist angles up to 16°.

The adoption of liquid hydrogen (LH₂) as a sustainable aviation fuel presents unique challenges, particularly regarding fuel tank dynamics during flight. Sloshing-induced thermodynamic changes in cryogenic tanks can significantly impact pressure stability and fuel management systems. This study addresses the critical need to model and understand these complex interactions to ensure safe and efficient LH₂ storage in dynamic flight environments.

An integrated computational framework was developed using Cranfield University's in-house thermal-fluid code as the foundation. Ludwig's thermal diffusivity model, originally validated for liquid nitrogen (LN₂), was adapted from a temperature-based to an energy-based approach and incorporated into the thermal-fluid model to capture sloshing dynamics. The enhanced model underwent validation against experimental LN₂ data before being applied to LH₂ systems under the assumption of similar cryogenic fluid behaviour. A comprehensive three-phase methodology encompassed validation, LH₂ case studies, and parametric analysis examining initial pressure, fill levels, and excitation characteristics.

Model validation achieved agreement with experimental data, showing maximum deviation of only 7%. LH₂ simulations revealed a sloshing-induced pressure drop of approximately 100 kPa over 40 seconds, attributed to enhanced condensation and turbulence at the liquid–vapour interface. Parametric analysis demonstrated that higher initial pressures and fill levels (85%) increased pressure drop rates by 5-8%. Notably, chaotic sloshing patterns produced substantially faster pressure decrease (5.2 kPa/s) compared to planar wave sloshing (3.2 kPa/s).

This research demonstrates the capability of the enhanced thermal-fluid model to simulate complex cryogenic tank behaviour under dynamic conditions. The findings provide valuable design insights for optimizing LH₂ tank configurations and operational strategies in future hydrogen-powered aircraft, contributing to the advancement of sustainable aviation technologies.

Aerospace engineering applications for air and space solutions have been becoming increasingly complex. They are required to provide high performance while keeping sustainable safety in place, particularly when human beings are involved in their operations. Multidisciplinary development teams of engineers, system architects, and other stakeholders typically work together to design such sophisticated man-made systems. They use computer models that have different views of the same aerospace system under development. However, technical collaboration between interested parties becomes a deadlock when representations from each collaborator must be cross-checked across multiple models (from other developers) to assess the impact of diverse viewpoints, e.g., how changes in the software design models affect control design models. Later changes in modelling augment risks and costs as they require synchronization of blueprints by updating, and then re-verifying as well as re-validating each of the model representations.

This paper presents details of the design process of a planet rover in which diverse stakeholders deal with aspects of the space ground vehicles. It includes preliminary results from requirements analysis and system design that are used to interlink system models to set an autonomous crosschecked-model design process. The approach is carried out by a multidisciplinary design framework for development of aerospace systems that is meant to reduce collaborative design efforts by enabling multiple developers to minimize potential design inconsistencies (that ultimately produce downtimes) by means of interconnecting their system models. The framework methodology is based on an Artificial Intelligence (AI) support to autonomously link the parametrical notations from different models for modelling analytics, and for further advice on dependencies between models and actions to be taken. Concluding remarks and future research are also presented.

Multirotor unmanned aerial vehicles increasingly require design optimisation balancing aerodynamic efficiency, noise emissions, and operational constraints. While machine learning surrogates enable rapid performance prediction, their opacity limits physical insight into how design variables, such as rotor speed, blade pitch, and geometry, interact to shape system behaviour. For safety-critical applications, this lack of transparency complicates certification and informed design decisions. This work presents a framework combining High Dimensional Model Representation (HDMR) and Kolmogorov–Arnold Modelling (KAM) to extract interpretable structure from rotor aerodynamic performance models.

HDMR provides variance-based sensitivity indices and quantifies parameter interactions through additive decomposition. KAM complements this with a compositional representation that identifies localised functional regimes and reveals how variable importance shifts across the design space. The methodology is first validated on analytical benchmark functions exhibiting tuneable interaction structures, then applied to a parametric study of rotor thrust and efficiency generated using a mid-fidelity vortex particle method. The framework provides a foundation for extending the analysis to multi-rotor configurations where parameter interactions become increasingly complex.

The results will demonstrate how HDMR effectively ranks dominant parameters globally, while KAM uncovers regime transitions corresponding to shifts in dominant parameter interactions across the design space. The combined approach supports transparent design exploration and provides a starting point for regime-aware optimisation strategies in multirotor development.

CubeSat development has historically been mission-specific, with each spacecraft built as a unique integration of avionics, structure, EPS, and ADCS hardware. While this enables tailored optimization, it also increases non-recurring engineering effort, extends integration schedules, and complicates qualification and acceptance testing for every new flight article. As small satellite missions expand beyond academic demonstration toward sustained science and commercial services, the lack of a standardized, reusable bus architecture has become a limiting factor for program throughput and cost efficiency. This work presents a 3U common CubeSat bus designed to interface directly with a 3U payload, forming a complete 6U spacecraft without requiring new bus-level avionics layout, system harnessing, or mechanical redesign for each mission.

The architecture establishes consistent mechanical mounting features, power distribution topology, thermal sink paths, compute resources, and safe-to-mate inhibit logic, enabling payloads to integrate through defined electrical and structural interfaces rather than bespoke bus adaptation. Ten 6U missions launched between 2018 and 2025 were evaluated to derive power, pointing, mass, and data-rate envelopes that guided sizing of solar generation, battery capacity, OBC throughput, and ADCS accommodation. Results indicate that a single, openly defined 3U bus can support multiple payload classes including imaging, RF communications, atmospheric science, and technology demonstration with minimal configuration changes.

This paper argues that the primary value of the proposed bus is not reliability as an outcome, but architectural standardization as the mechanism enabling repeatable manufacturing, faster ATP flow, and reduced development friction across mission sets.

Introduction: The red-tailed hawk (RTH) is a remarkable soaring raptor with a broad wingspan that grants it impressive aerodynamic efficiency, allowing for effortless gliding and rapid transitions into high-speed dives for instant prey capture. These capabilities make the RTH an ideal biological model for developing a feathered, morphing drone that is capable of achieving glider-like aerodynamic efficiency while retaining the maneuverability of quadcopters. Methods: This work presents “SMART Hawk,” a biomimetic, non-flapping Unmanned Aerial Vehicle (UAV) with wing- and tail-morphing capabilities. SMART Hawk’s target weight and size are based on the physical characteristics of a female RTH. A mathematical model was first developed in MachUpX to guide the selection of the design parameters for optimal aerodynamic performance at various morphing configurations. SMART Hawk’s wing incorporates artificial composite feathers that are distributed along the wingspan and connected to a linkage mechanism modeled after avian wing bones, allowing the wing to sweep in a way that mimics the natural deformation of the hawk’s wing natural. The wing’s airfoil profile was extracted from the cross-section of the RTH obtained through 3D scanning of an RTH cadaver. A compact tail mechanism was designed to actuate three degrees of freedom: pitching, tilting, and feather expansion. Composite shells, balsa wood, and 3D-printed ASA plastic were used strategically for the structural and load-bearing components. Propulsion was achieved via a single electric motor located at the nose of the fuselage. Computational fluid dynamics (CFD) and finite element analysis (FEA) simulations were performed to validate the design of all UAV components. Results: A proof-of-concept prototype was built and flight tests were performed to prove the effectiveness of the proposed design. Conclusions: The proposed biomimetic morphing UAV design can replicate the aerodynamic qualities of the RTH. The selected materials and servomotors enabled the achievement of the design objectives.

Hybrid rocket propulsion systems, utilizing liquid oxidizers in conjunction with solid hydrocarbon fuels, are recognized for their safety, environmental compatibility, operational simplicity, and capacity for precise control over thrust and impulse profiles. Thus, this emerging technology offers a promising alternative for future space propulsion applications. In the last decade, the Utah State University Propulsion Research Laboratory (PRL) has developed hybrid propulsion technologies for a diverse array of space mission applications. This paper summarizes key progress in the PRL’s green-hybrid propulsion research, offers insights for scalable design future strategies, and presents a road map for future applications.

At the PRL, significant advancements in hybrid rocket technology have addressed both performance and sustainability requirements. Among the key technological advances are a patented low-energy arc-ignition system, which ensures reliable and consistent ignition, and a digital throttle control that enables precise and deep throttling. These advancements significantly enhance flexibility in missions requiring orbital transfer, attitude control, and rendezvous operations. These advances enable hybrid propulsion systems to serve as viable competitors to conventional chemical and electric propulsion methods for space applications.

PRL research also focuses on environmental sustainability alongside ignition and throttling improvements. A key objective is to substitute toxic and hazardous hydrazine-based propellants with safer, environmentally friendly options for in-space propulsion. By leveraging modern additive manufacturing methods, PRL research has enabled fast, affordable solid-fuel production by using recycled and bio-based plastics. These advances have optimized fuels, resulting in lowered production costs, minimized environmental impact, and increased motor efficiency. These advanced fuel blends exhibit advantageous regression rates, consistent mechanical strength, and compatibility with current ignition and oxidizer-feed systems. This research shows that sustainable materials can be used in propulsion without compromising performance, enabling hybrid rocket engines to be competitive in this market space.

Antarctic ice-shelf movements are essential for controlling sea-level rise and global climate patterns. Continuous monitoring of the region poses substantial challenges due to the vast spatial extent, remote location, and harsh environmental conditions of Antarctica. The Antarctic Ice Exploration Experiment (AICE) mission concept is designed to address these challenges by tracking changes in Antarctic ice shelves using a satellite positioned in a highly elliptical orbit. The mission employs a Synthetic Aperture Radar (SAR)- equipped 12U-class small satellite to effectively monitor designated areas of interest (AOI) over a nominal two-year mission period.

The proposed mission utilizes a single-satellite-single-maneuver (SSSM) configuration to reduce the overall mass of the satellite while minimizing the Δv budget, thereby maintaining the possibility of accessing a high-inclination orbit without any plane-change maneuvers, thus maximizing polar region observation capability. Synthetic Aperture Radar (SAR) was selected as the primary scientific payload for the satellite to collect measurements of ice shelves for its ability to deliver high-resolution imagery in all-weather, day-and-night conditions, ensuring reliable coverage of Antarctic ice shelves. The radar payload features a 30° field of view to maximize the swath exceeding 400km during each pass.

To extend the observation time, the spacecraft is placed on a Molniya orbit with an inclination of 63.4°. It provides over 75% of the mission lifetime worth of observation window dedicated to Antarctic observations. The AICE mission demonstrates the feasibility of a single-satellite, single-maneuver spacecraft with SAR-enabled imagery, enabling sustained monitoring of the Antarctic region in highly elliptical orbits. The approach highlights a scalable and efficient solution for earth observation and information-gathering missions in long-term climate modeling, polar science, and the understanding of ice-sheet dynamics.

Introduction: Dust generation during lunar regolith excavation presents a major challenge for sustained surface operations, with implications for mechanical wear, optical degradation, thermal control, and crew safety. Most current excavation approaches address dust through mitigation techniques applied after particle liberation. In this work, we investigate an alternative strategy focused on dust prevention by design, using a staged excavation architecture intended to limit particle release at the source during material interaction.

System Design: We present the design of a dual-stage excavation and containment subsystem developed as the core functional element of a small-scale experimental platform. The architecture integrates material engagement, controlled internal transport, and temporary storage within a single enclosed cylindrical volume. This configuration is intended to minimize exposure of disturbed regolith to the surrounding environment during insertion and excavation, thereby reducing opportunities for particle entrainment and dispersion.

Experimental Methods: Preliminary experimental testing was conducted in open-air conditions using LMS-1E lunar regolith simulant. The excavation subsystem was mounted on a fixed test jig to isolate excavation behavior from vehicle dynamics, vibration, and wheel–terrain interactions. Dust dispersion was evaluated using gravity-collected particles captured on adhesive sampling pads placed at fixed locations around the excavation zone. Following each test, sampling pads were photographed and weighed to estimate relative dust deposition patterns. Ambient humidity was not controlled during testing.

Results and Discussion: Observations indicate reduced visible dust during both insertion and excavation phases, with dust deposition patterns suggesting localized particle release and limited spatial dispersion. Collected particles on sampling pads were concentrated near the excavation interface, consistent with the staged excavation hypothesis. While the results are preliminary and qualitative, they support the potential of staged excavation architecture as a viable approach for reducing dust generation during lunar regolith operations.

Space and rocket-related projects are notoriously expensive, with costs that often exceed initial budgets, especially for government-funded projects. These high costs pose a substantial barrier to expanding the capabilities of space exploration, satellite deployment, and related scientific research. To overcome these challenges, this study presents the optimization of a two-stage rocket for low Earth orbit (LEO) at a maximum altitude of 450 km and a 300 kg payload, focusing on aerodynamic and propulsion optimization to enhance fuel efficiency.

A multidisciplinary framework was applied, incorporating propulsion system analysis, and thermodynamic properties of the rocket nozzles using NASA’s Chemical Equilibrium with Applications (CEA) code. The Computational Fluid Dynamics (CFD) simulations were conducted in SolidWorks to refine the rocket's aerodynamic profile by optimizing fin geometries. The nozzle geometry was refined using Rao’s method, with MATLAB calculations applied to determine the optimal nozzle expansion ratio. NASA’s General Mission Analysis Tool (GMAT) was utilized to perform trajectory optimization and propellant mass calculations. The nozzle performance was validated through benchmarking against comparable-class small-lift launch vehicles.

The optimized launch vehicle achieves a total lift-off mass of 13,000 kg, including 11,200 kg of propellant and a 300 kg payload. The first stage produces 235 kN of thrust with a burn time of 155 s and a specific impulse of 265 s, while the second stage delivers 25 kN of thrust over 350 s with a specific impulse of 340 s. The combined two-stage propulsion system provides a total ΔV of 13.71 km/s, enabling a final orbital altitude of approximately 400 km. These findings contribute to more sustainable two-stage propulsion systems by enhancing fuel efficiency and sustainability in small-lift launch vehicles.