Advanced European Re-Entry System Based on Inflatable Heat Shields: Detailed Design (EFESTO project)

The European Union H2020 EFESTO project is coordinated by DEIMOS Space with the end goals of improving the European TRL of Inflatable Heat Shields for re-entry vehicles (from 3 to 4/5) and paving the way towards further improvements (TRL 6 with a future In-Orbit Demonstrator, IOD). This paper presents the project objectives and the initial results of the detailed design of atmospheric entry missions based on the applications of advanced thermal protection systems implementing inflatable heat shields (flexible TPS and inflatable structures), according to aerothermodynamics constraints for future in-orbit demonstration. Placing the future IOD mission in the context of ongoing and future efforts in the European context is also one of the project goals. Two key applications, Mars Robotic Exploration and Reusable Small Launchers Upper Stages, have been identified. For the Mars Application, the robotic exploration mission class resulted in a 10 m diameter Hypersonic Inflatable Aerodynamic Decelerator (HIAD) class, combined with Supersonic Retro-Propulsion (SRP, activated about Mach 2.3) to deliver about 2800 kg of payload at MOLA +2 km. For the Earth Application, the VEGA upper stage (AVUM) has been selected as baseline case study. The current mission foresees a deorbiting from SSO orbit, a controlled entry phase (BC of about 30 kg/m2) and combines the use of a HIAD (4.5m diameter class) with parachutes and parafoil for Mid-Air-Capturing (MAR) with a helicopter. Beyond feasibility of the entry mission phase and system design with an inflated IAD, integration aspects have a key impact in the specific design solutions adopted, due to the nature of an inflatable heatshield. For both considered application cases feasible architectures are developed responding to the challenge of integrating the HIAD into the system in compliance with geometric and functional requirements. While the HIAD in folded state prior to inflation must fit in the available volume, it has limitations with respect to the density imposing a minimum cross section of the stowage volume. Simultaneously requirements with respect to the centre of gravity position during re-entry with an inflated HIAD must be respected for stability and controllability reasons. Other architectural considerations such as payload integration for the application on a launcher upper stage must be considered. Finally, heat loads constraints are considered for the trajectory and TPS deign choices due to important fluid-structure interactions. This project has received funding from the European Union’s Horizon 2020 research and innovation programme under grant agreement No 821801

EURASTROS ascent trajectory and abort analysis

The EURASTROS study [1] was a joint study between Ariane Group GmbH and the German Aerospace Center (DLR), exploring astronautic transport capabilities of Ariane 6. The study included the preliminary design of a crew module (CM) [2], a service module (SM) [3], a performance analysis of possible Launch Abort System (LAS) concepts and a cost analysis [4]. This paper presents the work performed with the purpose to define the general ascent strategy and to design suitable end-to-end abort strategies compatible with the mission and system requirements. It shows the selected reference ascent trajectory and covers detailed preliminary analyses of the exo- and endo-atmospheric abort scenarios regarded within the EURASTROS project. All analyses were closely tied to continuous optimization of the Ariane 6 ascent trajectory and corresponding iteration of achievable payload performances, using astronautic transport to an ISS orbit as reference study case. The conducted analyses explored multiple possibilities for performance adjustment in agreement with human rated mission requirements, while also respecting space debris mitigation standards. Eventually these investigations concluded in the proposal of 11 Mg propellant un-loading of the Upper Liquid Propulsion Module (ULPM) of the Ariane 64 (A64) with respect to standard GTO upper stage fuel, yielding a trajectory enabling for maximum performance and also sufficient safety. The crew module, envisioned for a total crew of up to three astronauts, was designed with close resemblance to the Apollo CM, with a Service Module (SM) for exo-atmospheric flight at the rear and a Launch Abort System (LAS) for endo-atmospheric abort at the front, but with volume reduced by 28 %. For the LAS design, the most promising configurations were designed towards minimum thrust allowing to reach a minimum safety distance of 200 m to the launch vehicle within 3.5 seconds after separation and to fulfill further safety requirements in case of an abort from launch pad. Simulations conducted at DLR concluded that the bare minimum thrust should exceed 950 kN for both concepts in order to meet imposed safety requirements. Based on these results a subsequent mass budget estimation yielded a total mass estimate between 4700 – 5400 kg for the LAS. The SM is based on the ASTRIS kick stage and utilizes the engine BERTA running on storable fuel. This module is dimensioned to provide continuous exo-atmospheric launch abort capabilities while preventing any potential impact on populated areas on European and Eurasian soil. For such an abort scenario, the Last Direct Re-entry Point (LDR) after which an abort-to-orbit would be executed, has been defined. Influences of the corresponding maneuver pitching angles on ballistic downrange as well as on required time for abort-to-orbit are explored and the SM minimum thrust was adjusted to yield sufficient performance capabilities. Eventually, six abort modes for exo-atmospheric flight are proposed and discussed

Assessment of VTVL and VTHL Reusable First Stages

Two-stage vertical take-off vertical landing (VTVL) and vertical take-off horizontal landing (VTHL) partially reusable launcher configurations are systematically analyzed. The investigated configurations consider a reusable first stage that either performs a landing at the launch site (return to launch site RTLS) or a landing downrange of the launch site (downrange landing DRL). The considered propellant combinations include LOX/LH2, LOX/LCH4 and LOX/RP-1. Configurations based on staged combustion and gas generator cycle engines are analyzed. The same engines however with different expansion ratios are used on the reusable first stages and the expendable upper stages. Special emphasis is put on analyzing the different configurations under similar design assumptions that allow a comparison of gross lift-off masses, stage lift-off masses, stage structural indices as well as loads encountered by the reusable stages during atmospheric reentry. Based on this comparison benefits and drawbacks of the investigated RLV configurations are discussed

Family of Launchers Approach vs. “Big-Size-Fits-All”

One option for future space transportation concepts could be a family of launchers supporting a wide range of payload performance. The idea bases upon using “building blocks” of common stages or main propulsion rocket engines and applying them in a modular way. A somehow contrarious option is a single TSTO launch vehicle serving all kinds of missions, even those which have payload mass requirements much below the design capacity. The technical investigations described in this paper evaluate the two antipodal design approaches of either establishing a launcher family consisting of modular building blocks or choosing a full-size launcher which serves all missions with minimal adaptations of the upper- and kick-stage selection. The paper summarizes major results of the preliminary technical design process. The overall shape and aerodynamic configuration, the propulsion and feed system, the architecture of the stages are described and different technical solutions are compared. Payload performance is optimized for the different concepts in the GTO-mission, manned flight to ISS and to SSO. The winged configurations’ controllability in hypersonic reentry and subsequent subsonic flight is assessed. The study is completed by a relative comparison of to be expected RC/NRC of the different launcher concepts

ANALYSIS OF VTVL AND VTHL REUSABLE LAUNCH VEHICLE CONFIGURATIONS

The German Aerospace Centre (DLR) is conducting systematic analyses of reusable space transportation configurations. Two-stage vertical take-off vertical landing (VTVL) and winged, vertical take-off horizontal landing (VTHL) partially reusable launcher configurations are systematically analyzed. The investigated configurations consider reusable first stages that either perform a return to launch site or land down range of the launch site. The propellant combinations analyzed include LOX/LH2, LOX/LCH4 and LOX/RP-1. Staged combustion and gas generator cycle engines are taken into account. The same type of engines with different expansion ratios are used on the reusable first stages and the expendable upper stages. Major analysis objectives are the comparison of various reusable launch vehicle configurations under similar design assumptions as well as the identification of their critical aspects, benefits and drawbacks

Development of Inflatable Heat Shield Technology for Re-Entry Systems in EFESTO project

The EFESTO project is funded by the European Union H2020 program. The purpose is to increase European capabilities in designing Inflatable Heat Shields for re-entry vehicles. The technology of inflatable heat shields enables increasing the spectrum of space-based applications as it provides effective heat protection and deceleration capabilities for atmospheric descent while being comparatively mass and volume efficient which is a significant asset for a space mission. The use of inflatable heat shields for Mars exploration and for Earth re-entry of a launcher upper-stage for later reuse were selected at the initial study phase as potential application for the HIAD technology. These two application cases are to demonstrate the performance of this technology under realistic conditions and to provide a representative study frame for the design of inflatable heat shields trained at tangible applications. In the first part of the project, the work focused on the system design of both study cases. This work yielded an inflatable heat shield design that shows a reduced complexity in geometry compared to the initial design and is scalable for other applications. Several stacks of material layers for the Flexible Thermal Protection System (F-TPS) were traded against each other before selecting one reference definition for the consecutive project phases. An intense test activity followed this phase. Part of the tests served for verifying the thermal performances of the F-TPS under relevant aerothermal environment using the plasma wind tunnel test infrastructure available within the consortium. In addition, a high-fidelity inflatable structure ground demonstrator was manufactured. This demonstrator served to consolidate the mechanical characterization of the inflatable system. This testing activity provided the data used for numerical cross-correlation and experimental-numerical rebuilding. Eventually, computational folding analysis completed the numerical activity during this project phase. The final project phase is dedicated to the preliminary design of an in-orbit demonstration mission for the technology and design of the technology development roadmap. This potential future In-Orbit Demonstrator (IOD) shall provide knowledge of the system performance while evolving in a relevant environment. This will provide in-flight verification and validation of the developed inflatable heat shield technology. This paper gives an overview of the project with a focus on system aspects of the EFESTO project about to be completed in the coming weeks

Outlook on the New Generation of European Reusable Launchers

The technical investigations described in this paper evaluate the two seemingly antipodal design approaches of either establishing a launcher family consisting of modular building blocks or choosing a full-size reusable launcher stage which serves all missions with adaptations limited to the upper- and kick-stage selection. The paper summarizes major results of the preliminary technical design process iteratively performed at DLR-SART. The overall shape and aerodynamic configuration, the propulsion, the architectures of the stages are described and different technical solutions are compared. Payload performance is optimized for the different concepts in the GTO-mission, manned flight to ISS and to SSO. The winged configurations’ controllability in hypersonic reentry and subsequent subsonic flight is assessed

The EFESTO Project: Advanced European Re-Entry System based on Inflatable Heat Shield

EFESTO is a project funded by the European Union H2020 program aiming for a revamp and growth of European know-how and systems engineering capabilities in the strategic field of Inflatable Heat Shield technology for re-entry vehicles. This project analyzes the use of Inflatable Heat Shields for Mars exploration and Earth re-entry applications that served as representative study-cases. In addition to design activities at system and sub-system levels, the EFESTO team focused on testing the aerothermodynamic properties of the Flexible TPS and the mechanical characteristics of the shield, the latter exploiting a manufactured high-fidelity Inflatable Structure demonstrator. The data gathered from the two test campaigns additionally served for experimental-numerical rebuilding and cross-correlation. Finally, a phase-0 feasibility study defined a preliminary IOD mission design to enable in-flight verification and validation of the critical technologies. This paper will present the whole excursus of the project, including the key phases of use-cases survey and investigation, mission scenarios definition and analysis, system engineering and sub-system design, technology development and ground demonstration, future roadmap identification with reference IOD feasibility analysis and early definition. The project achievements have improved the European TRL of Inflatable Heat Shields from 3 to 4/5, thus paving the way towards further developments in the mid-term future. This project has received funding from the European Union's Horizon 2020 research and innovation program under grant agreement No 821801

EFESTO - advancing European hypersonic inflatable heatshield technology for Earth recovery and Mars high-mass delivery missions

The European Union H2020 EFESTO project is coordinated by DEIMOS Space with the end goals of improving the European TRL of Inflatable Heat Shields for re-entry vehicles (from 3 to 4/5) and paving the way towards further improvements (TRL 6 with a future In-Orbit Demonstrator). This paper presents the project objectives and provides with a general overview of the activities ongoing and planned for the next two years, promoting its position in the frame of a European re-entry technology roadmap. EFESTO aims at (1) the definition of critical space mission scenarios (Earth and Mars applications) enabled by the use of advanced inflatable Thermal Protection Systems (TPS), (2) characterization of the operative environment and (3) validation by tests of both the flexible materials needed for the thermal protection (flexible thermal blanket will be tested in arcjet facility in both Earth and Martian environments) and the inflatable structure at 1:2 scale (exploring the morphing dynamics and materials response from packed to fully inflated configuration). These results will be injected into the consolidated design of a future In-Orbit Demonstrator (IOD) mission

A viable and sustainable European path into space – for cargo and astronauts

A think tank of DLR space-research institutes and ArianeGroup reflected on the launcher perspectives, broadening the mission capabilities for European access to space including human mission towards LEO, supporting potential future commercial markets. The Ariane 6 launcher is to be evolved with perspective of enhanced versatility and reduced costs. Also, a new generation of modern, high-performance launchers is to be prepared that is capable of serving all major missions relevant to Europe. The paper summarizes major results of preliminary technical design studies performed in Germany, both for a future Ariane 6 evolution including astronaut transport to and from LEO and for the next generation of partially reusable concepts. Return of the RLV is pondering options of powered descent and vertical landing and smart winged return technologies with horizontal landing close to the launch site. The main propulsion system is considering LOX-LH2 in different cycle architectures as well as LOX-LCH4 of the ongoing PROMETHEUS development program. For the studied concepts, the overall shape and aerodynamic configuration, the propulsion system, the architecture and lay-out of the stages are described and different technical solutions of RLV concepts are compared