Research paperPhase-A design of a reusable re-entry vehicle
Introduction
Over the past few years, multi-functional space vehicles operating in different mission scenarios as independent platforms or as support vehicles (manned or unmanned) for the International Space Station (ISS), are becoming the upcoming standard of the future spacecraft design [1]. In this framework, several mandatory requirements are outlined by space agencies, eventually affecting the feasibility of a space programme. Future re-entry vehicles should be reusable with the capability to perform several flights without Thermal Protection System (TPS) refurbishment/replacement, as for the space Shuttle. Moreover, improved aerodynamic performances should enhance the possibility to land on commercial runways, thus allowing a comfortable landing environment (reduced g-loads) for non trained or injured crew transport [2].
All above demands are strongly related with each other, and reflect a multi-disciplinary characterization of re-entry vehicle design. Additionally, their trade-off lead to a very different aeroshape design if compared with the Space Shuttle one. System reusability results in lower costs to payload transportation once reliability of thermal protection materials is assured [3]. A reusable vehicle implies that the aero-thermal loading environment allows the adoption of an high performance TPS exclusively cooled by thermal radiation. Therefore, the spacecraft does not require refurbishment of TPS parts between one flight and another. Furthermore, a reusable TPS reduces the related cost to put payloads into orbit, and the obvious higher risk associated with more severe thermal loads.
In the above scenario, an aeroshape designed as a blended wing–body (BWB) with a delta-shaped planform, with and a relatively small ratio, can slowly decelerates at high altitude, thus increasing the flight time and experiencing a reduced convective heat flux. This concept vehicle should benefit of a short refurbishment time for a high turn-around mission rate. Additionally, the increased duration of flight improves the cross range performance of the glider, thus giving the possibility to customize the landing spot. As previously outlined, a modern spacecraft design has to satisfy a very challenging picture which relates safety, operability, structural and aerodynamic constraints, cost reduction, and finally flight comfort. Furthermore, several design choices that are strongly desirable in the high-speed regime, reveals poor performance at low-speed and even more at landing. Eventually, the design of new-generation vehicles is not only a multidisciplinary task, but it also requires the development of uncommon solutions in order to provide an answer to antithetical design requirements. Therefore, it is highly believable that a new-generation of space gliders can only pursue holistic design specifications (referenced to the entire mission) if the vehicle shape is derived from a seamless, multi-purpose design instead of the mere superposition of different singular-focused concepts. The Dream Chaser, jointly designed by (NASA and SNC) and featuring a lifting-body aeroshape with small wings, is one of the most interesting design candidate vehicle up to now. It allows gliding, banking and landing to assigned conventional runways due to its good aerodynamic performances. The Dream Chaser blunt nose-profile and flat bottomed body represents a notable example of an accurate synthesis of all aerodynamic features experienced during the re-entry. The morphological complexity of the Dream Chaser reflects the great effort to obtain a high-performance design which efficiently takes into account for a large number of design variables. This category of design procedures is formulated as multi-objective, multi-disciplinary and multi-level design optimization procedure. The design, in fact, involves several evaluations performed in sub-discipline analysis using different levels of accuracy. Multi-disciplinary Design Optimization (MDO) represents a powerful tool to derive shapes from the mutual contribution of different sub-discipline, as Aerodynamics, Aerothermodynamics, Flight Mechanics, and Preliminary Design [4], [5]. An important starting point for a successful optimization-based design is the shape parameterization. A shape generator, efficiently feeds a chain of sub-discipline analyses, and elaborate the design variables to an optimization algorithm often performed with evolutionary algorithm (i.e., Genetic Algorithm, GA).
In this framework, the paper deals with a detailed Phase-A Design Environment (DE) developed to obtain a suitable spacecraft candidate for a return mission from Low Earth Orbit (LEO), which ends with a landing on a horizontal conventional runway. The DE regroups all the necessary phases that drive to the definition of a suitable preliminary design configuration. It starts from a conceptual design obtained with a MDO procedure and concludes, after a feedback by several accurate Computational Fluid Dynamics (CFD) simulations. The definition of a feasible configuration is then performed by adding two independent body flaps, to provide control along both the longitudinal and lateral-directional stability axes. Moreover, the assessment of a LEO re-entry based on an optimized prescribed guidance law, is also performed considering thermal and aerodynamic performance objectives. This approach, based on the authors’ knowledge, summarizes in a unique procedure several methodologies with an increasing level of fidelity. The DE exploits an in-house one-piece shape generation procedure and the capability to determine the best guidance profile for a given re-entry mission. Both shape generation and guidance law are determined adopting GA algorithm.
Finally, it is worth noting that this work represents a step forward within a wider research activity by the authors concerning the development of a multi-purpose re-usable vehicle for a low cost access-to-space in LEO, and a safe ending with horizontal landing on a conventional runway [6], [7], [8].
Section snippets
Related work
Several studies that use a multidisciplinary-multilevel analysis to design re-entry vehicles are already published in literature. Tava et al. [9] addressed a multi-objective aeroshape optimization to maximize the cross range and minimize the total heat absorbed by the spacecraft. A simple capsule geometry was obtained by a solid rotation of a parametric curve defining the outline of the shape. Chiba et al. [10] modelled the wing–body section of a booster using an artificial neural network
Mission requirements and constraints
The concept aeroshape is designed for a LEO re-entry mission to and from the ISS and can be used either as emergency vehicle or crew return vehicle (from two to four astronauts), see Fig. 1. Initial re-entry altitude and speed are set equal to and , respectively. Spacecraft mass target is about .
Several mission constraints are accounted in the vehicle design, such as: maximum heat flux limit ; maximum dynamic pressure ; maximum normal load factor
Overview of design procedure
The design framework can be represented as a multi-level chain process, as shown in Fig. 2, where several stages, featuring increased levels of design refinement, are related through within a unique procedure. The multi-level framework outlines a bottom-up approach where simplified assumptions, used to handle with the complexity of the problem, are progressively removed. The framework comprises three phase; (i) conceptual design phase; (ii) Phase-A design; (iii) guidance optimization. The
Shape definition
The shape model which defines the concept configuration features the absence of explicit parametric support surfaces, as detailed in Ref. [19]. This peculiarity allows a seamless blending between body and wing, for each configuration. A generic shape instance is obtained from a three-dimensional parametric wireframe (see Fig. 5a–b), which is used to drive a self-organized cloud of points. Wire-frame parameters define the change of morphology. Additionally, points (see Fig. 5c–d) representing
Phase-A design: Aerodynamics and aerothermodynamics
The DE representing the conceptual configuration is assumed frozen in the following stages of design procedure. The next level of design chain in Fig. 2 is the Preliminary Design Phase. At first, validation and a refinement phase of aerodynamic computations is performed. The mission goal considered in this analysis is the returning of a crew from LEO, with landing on conventional runway at about , according to a Shuttle-like touchdown [27]. The spacecraft re-entry flight covers different
Re-entry flight and guidance optimization
The shuttle-like trajectory in the drag acceleration–velocity plane when flown without a bank angle modulation exhibits several oscillations during hypersonic phase which should avoided for safety reasons. Moreover, high values of drag accelerations are potentially harmful for TPS integrity due to consequent high thermal load oscillations. Therefore, a re-entry strategy which envisages a longer duration re-entry, and consequently a gradual conversion of spacecraft’s potential energy into
Conclusions
This paper dealt with a detailed design procedure developed to obtain a suitable candidate for a LEO return mission, which ends with a landing on an horizontal conventional runway. Several aspect of design phase were addressed in a single procedure which regroups the necessary phases for a preliminary design study. Specifically, a multi-objective optimization procedure was used to find into a wide search space, several antithetical design choices. Furthermore, it was shown that the peculiar
Declaration of Competing Interest
The authors declare that they have no known competing financial interests or personal relationships that could have appeared to influence the work reported in this paper.
Prof Giuseppe Pezzella is lecturer in Aircraft Aerodynamics and Aerothermodynamics at the Department of Engineering of the University of Campania, Italy. He was also research engineer at CIRA (Italian Aerospace Research Centre) in the field of hypersonic aerodynamics and aerothermodynamics. He achieved the degree in Aerospace Engineering on 1999 and the Ph.D. degree in Aerospace Engineering on 2004 at the University of Naples “Federico II”. He was involved in several re-entry vehicle design
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Prof Giuseppe Pezzella is lecturer in Aircraft Aerodynamics and Aerothermodynamics at the Department of Engineering of the University of Campania, Italy. He was also research engineer at CIRA (Italian Aerospace Research Centre) in the field of hypersonic aerodynamics and aerothermodynamics. He achieved the degree in Aerospace Engineering on 1999 and the Ph.D. degree in Aerospace Engineering on 2004 at the University of Naples “Federico II”. He was involved in several re-entry vehicle design activities, in particular in the frame of ESA (EXPERT, FLPP, HIGH-LIFT, RASTAS-SPEAR, SCRAMSPACE, LAPCAT-II, IXV, HEXAFLY, HEXAFLY-INT) and National (PRO.R.A. USV) programmes.