Reusable Launch Vehicle Concept Research Articles

Overview of system study on recovery methods for reusable first stages of future European launchers

AbstractThe design of a reusable launch vehicle implies the need to provide for a means to safely retrieve the component to be re-used. Following economic considerations, reusable launch vehicle concepts tend to be designed such that large parts, like entire stages, are to be recovered. These are usually significant in size and weight and have acquired a considerable amount of energy during their primary mission. This poses the challenge of how to recover them in a way that makes it available for further re-uses. In the past and present, different methods were and are used. Depending on the selected recovery method, the system design is very different necessitating different technologies and competencies to be acquired for a successful design. Two major classes of recovery methods can be distinguished: those recovery methods ending with a vertical landing of the reusable stage and those ending with a horizontal landing. Both have their own benefits and drawbacks. In 2016, The German Aerospace Centre DLR has initiated a large in-house study with the aim of investigating, in a comparative manner on system level, both classes of recovery methods on a system level for two-stage-to-orbit launch vehicles with a reusable first stage and an expendable upper stage to be operated within a European context. Fuel choice and engine cycle were major design parameters that were considered during the study. The present paper presents the framework of this study describing the adopted study logic, providing an overview of the major findings obtained at the end of the first study phase and gives an outlook to the work of the second study phase. It ends with providing a view of a possible demonstrator and technology roadmap toward the realization of an operational two-stage-to-orbit launch system with a reusable first stage.

Advanced modeling and trajectory optimization of the in-air-capturing maneuver for winged RLVs

Future reusable launch vehicle concepts and their key technologies have been investigated within the DLR research project AKIRA. In this context, several return options for reusable launch vehicles (RLV) were categorized by vertical (SpaceX, Blue Origin) or horizontal landing strategies (Space Shuttle), and then systematically evaluated based on their influence on overall design and technical feasibility. In general, system dynamics, guidance, and control aspects are of special importance within preliminary design studies, in particular if complex and difficult maneuvers like the DLR-patented in-air-capturing method are considered. In this case, the unpowered winged RLV is captured during descent by an aerodynamically controlled capturing device which is connected to an aircraft by a cable. After successful capturing, the launch vehicle is towed back to its landing site.In previous studies, the technical feasibility of the in-air-capturing maneuver was mainly assessed by simulations for an aerodynamically controlled RLV and an aircraft which is assumed to be passive. In contrast to this, we consider an optimal control approach to the problem of in-air-capturing, investigating both passive and active (cooperative) RLV and aircraft operations. To study the risk of failure of the in-air-capturing maneuver, both the initial capturing approach and a subsequent second attempt for recapture after an initial miss are analyzed. For this purpose, a multi-disciplinary multibody modeling and simulation framework based on the object-oriented modeling language MODELICA is used for the consistent flight dynamics modeling of each vehicle including a rigid cable connecting the aircraft and its capturing device. The trajectory optimization results provide an overview of the dynamic behavior of the multibody system for several constraints and flight conditions. Additionally, the results show that for a successful in-air-capturing maneuver with minimum control effort and multiple recapturing attempts, an actively controlled aircraft with drag-increasing subsystems and a cooperative launch vehicle maintaining a suitable flight path angle are recommended. The obtained reference trajectories can be used for future controllability studies and control system design considering a flexible cable and disturbances.

Multidisciplinary modeling and simulation framework for launch vehicle system dynamics and control

Future concepts and key technologies for reusable launch vehicles are currently investigated by the DLR project AKIRA, focusing on vertical takeoff and horizontal landing (VTHL), as well as horizontal takeoff and horizontal landing (HTHL) concepts. Dedicated developments of multidisciplinary frameworks for launch vehicle modeling and preliminary design optimization have been presented in the relevant literature. These activities are often performed by several independent and discipline-specific tools; such an approach can only account for limited interactions of the involved disciplines with the overall system dynamics.Therefore, the objective of this paper is to focus on a multidisciplinary launch vehicle dynamics modeling, guidance, and control framework to support reusable launch vehicle design activities at DLR while taking into account the highly interconnected disciplines involved and changing environmental conditions. The modeling framework is based on the object-oriented, multidisciplinary, and equation-based modeling language MODELICA. Dedicated 3-DOF and 6-DOF model implementations, covering the kinematics and dynamics formulation, environmental effects, aerodynamics, and propulsion models are presented.Within this framework, a method to obtain a direct connection between 3-DOF and 6-DOF models is shown. This is done by considering results from the trajectory optimization package ‘trajOpt’ in combination with nonlinear 6-DOF inverse models obtained automatically by MODELICA. Angular rates and resulting moments can be retrieved by this intermediate 6-DOF modeling approach for subsequent controllability studies. We discuss some of these benefits in terms of nonlinear flight control simulations for an HTHL reusable launch vehicle concept.

Aerodynamic and aerothermodynamic design of Future Launchers Preparatory Program concepts

In the frame of the Future Launchers Preparatory Program, being developed by the European Space Agency, several Reusable Launch Vehicle concepts are under investigation. In this paper the current design activities at phase-A level, carried out by Centro Italiano Ricerche Aerospaziali, are described. The goal has been to define the preliminary hypersonic aerodynamic and aerothermodynamic databases of each vehicle for concept competition. To this end, different design approaches have been addressed as engineering methods and computational fluid dynamics. The re-entry scenario with the corresponding loading environment for each launcher concept is reported and analyzed. The final results, applicable for the prosecution of the launcher design activities, are that, at the condition of descent trajectory peak heating, the vehicle features a stagnation point heat flux ranging form about 500 to 40 kW/m2 whereas the aerodynamic lift-to-drag ratio at 40 deg angle of attach is about 1.0.

RLV candidates for European Future Launchers Preparatory Programme

ESA's Future Launcher Preparatory Programme (FLPP) includes system activities and technology developments on materials and structures and propulsion. The overall approach is system-driven, i.e. FLPP aims at defining reference launcher concepts from which will be derived a demonstration logic including in-flight experimental vehicles, and technology development requirements. The paper addresses the launcher concepts studies, and more specifically the analysis of Reusable Launch Vehicle (RLV) concepts that has been performed in the first part of the FLPP program. The paper provides first elements on the mission requirements, the operations and the ground facilities for the RLV concepts selected as potential FLPP candidates. In particular, operations will likely have a heavy impact on the operability and on the launch cost of RLVs; in this sense, most of the paper content is dedicated to discuss specific operation features, and related infrastructure elements, linked to the considered RLVs. At last, a brief presentation of the next steps of the launcher system activities inside the FLPP is given.

Data Analysis of Phoenix Reusable Launch Vehicle Demonstrator Flight Test

The Phoenix vehicle was designed to flight demonstrate the automatic and unpowered horizontal landing of a representative, winged reusable launch vehicle. The shape of the test vehicle was derived from the suborbital reusable launch vehicle concept Hopper. Three automatic landing tests were completed successfully in May 2004. Methods of system identification were applied to the flight data to evaluate the performance and to improve the design models and databases for future applications. A specific emphasis was placed on the evaluation of the onboard navigation system, air data sensor, aerodynamic model, landing gear effects and ground-roll characteristics. This paper gives a brief overview of the Phoenix mission and elaborates on the flight data analysis and of the preceding wind tunnel campaigns, to allow a comparison of results from different approaches.

Representation Language for Evaluating Reusable Launch Vehicle Concepts

†Evaluation of launch vehicle concepts requires analysis by experts in vehicle sizing, trajectory, propulsion, subsystems, operations, cost, and reliability. A common data model is required to ensure consistent analysis across the various disciplines. This paper describes the development of the Launch Vehicle Language to support evaluation of Second Generation launch vehicle concepts. Vocabulary definition, development of data structures, and implementation in XML are discussed. Future enhancements to facilitate interdisciplinary data exchange are described.

Comparison of Glideback and Flyback Boosters

The aim of the present study is to investigate alternative reusable launch vehicle concepts and to provide a comparison of two different vehicle cone gurations. The e rst concept vehicle uses airbreathing engines to perform a powered booster e yback mission, whereas the second concept employs only aerodynamic lift to achieve booster e yback. The optimization software employs a system parameter optimization, in conjunction with a trajectory parameter optimization, to calculate the optimal mass split and corresponding optimal trajectory. A reference mission is used to comparelaunch vehicle payloads fora e xed grossliftoffweight, assuming launch from Woomera Prohibited Area, South Australia. The payload for a vehicle using a powered e yback strategy is found to be considerably higher than that obtained by a vehicle employing an aerodynamic glide return e ight.

Eigenstructure vs Constrained H Design for Hypersonic Winged Cone

HE purpose of this study is to highlight some specie c decoupling issues in longitudinal control of the newer generation of airbreathing launch vehicles in hypersonic regime. The so-called winged cone, a reusable launch vehicle concept still being considered by NASA, is chosen as benchmark model. Critical in longitudinal control is the placement of the closed-loop poles at locations specie ed by, for example, the military standard e ying qualities of piloted aircraft (MILSPEC), along with the achievement of a certain degree of decoupling between the phugoid and the short-period modes. More specie cally, because airbreathing vehicles are vulnerable to change in angle of attack, itis imperative to achieve accurate angle-of-attack control and decouple it from the phugoid mode. This decoupling is traditionally achieved via the Shapiro eigenstructureassignment.Inthispaper,wedevelopanalternateapproach basedontranscribingtheamountofundesirablecouplingintermsof the H 1 normofsomespecie c closed-looptransfermatrixand using genetic algorithms to achieve the least possible H 1 norm subject

Payload Capability of Kerosene-Fueled Mini Shuttle

Introduction T HE Space Shuttle or Space Transportation System (STS) is a beautifully designed reusable launch vehicle (RLV). With the exception of the external tank (ET), all of its components are recovered, refurbished, and used again. Despite its many critics and high operating costs, it has been a workhorse in space for nearly 15 years of service with an unparalleled record of safety and reliability. The one  ight in more than 75 that ended tragically was found to have resulted from an O-ring failure in one solid rocket booster (SRB), which could have been avoided with proper launch precautions. Now, however, NASA is intent on developing a replacement spacecraft for the Shuttle thatwill be available in about 10–15 years and will provide considerably cheaper access to space. In spite of the demise of the National Aerospace Plane (NASP) program, the concept of a reusable single-stage-to-orbit (SSTO) vehicle is still considered viable.2;3 As a result, in 1996 a three-year contract was awarded to Lockheed Martin for development of the X-33 technology demonstrator.4 This is a subscaleprototypeof the Lockheed Martin VentureStar,which is envisionedas an SSTO RLV to replace the Shuttle.4 In its comprehensive access-to-space study, NASA considered three alternative courses of action to meet anticipated needs of the United States in launch capability through the year 2030. 1) Make necessary upgrades and continue primary reliance on the Space Shuttle and today’s  eet of expendable launch vehicles (ELVs) through 2030. 2) Develop a new expendable launch system that utilizes today’s state-of-the-art technology and begin the transition from the Shuttle and current ELVs in 2005. 3) Develop a new reusable advanced-technology,next-generation launch system and make the transition from today’s Shuttle and ELVs from 2006 to 2010. In the Ž nal analysis, option 3 was deemed the most desirable and was recommended for consideration by the President’s OfŽ ce of Science and Technology Policy. There are three alternative architectures in option 3 consisting of two reusable SSTO vehicles and one reusable two-stage vehicle. Two of the vehiclesuse air-breathing/rocket propulsionand one uses only rocket propulsion. Since publication of the NASA study in 1994, several option-3-type design concepts using only rocket propulsion have been reported. First, there is the NASA reference winged-body, vertical takeoff/horizontal landing (VTHL) design, which is a combination of a long cylindrical fuselage and a delta wing. Second, there is a cone-shaped,vertical takeoff/vertical landing vehicle without wings.6 Third, there is the delta-wing-shaped, lifting-body VTHL design, which has evolved into the Lockheed Martin X-33/VentureStar vehicle.4 A fourth reusable SSTO design is the wingless bent-biconic vehicle, which relies on a drag chute for stabilizationduring transonic  ight and a para-wing for landing. The authors of the paper on this last design point out some stability and control problems with the Ž rst two designs and a structural weight problemwith the lifting-bodydesign of the LockheedMartin VentureStar.8 Perhaps it might be well for NASA to reconsider the viability of the SSTO RLV concept. The NASP experience was a costly and wasteful one, of both time and money, and should not be repeated. Moreover, there is a new entry into the RLV sweepstakes that just might, when combined with the Shuttle-II concept, be the answer