In the last decade, the rapid and successful development of reusable launch systems such as SpaceX’ Falcon 9 demonstrated both the operational feasibility of reusable launchers and their economic viability. The objective of recovering a launcher or a launcher’s booster requires to safely return the launch vehicle from orbital or sub-orbital conditions to a soft landing. To increase the reusability, decrease the turnaround time and reduce costs, a precise touchdown on a pre-defined landing site or on a floating barge on the Ocean is preferred to splashdown in the water, due mainly to the highly detrimental effect of the salted water on the launcher components and equipment. The project RETALT (Retro Propulsion Assisted Landing Technologies) was funded by the EU Horizon 2020 program to study and develop critical technologies for launcher recovery based on retro-propulsion. In this context, and based on in-house experience and tools, DEIMOS Space carried out the mission engineering of the RETALT1 vehicle concept to assess the feasibility of a return mission, from a wide range of Main Engine Cut-Off (MECO) conditions, when the stage is separated from the rest of the launch vehicle, in line with the available propellant budget, and while maintaining the peak entry conditions within acceptable limits. Either a DownRange Landing (DRL) on a drone ship at sea or a Return To Launch Site (RTLS) to land in the proximity of the launch pad is performed based on the velocity and distance at MECO from the launch site. For the landing burn, a safe splashdown approach has been implemented to avoid damaging the ground infrastructure in case of anomalies during the flight. Based on the mission feasibility assessment, the needs for the vehicle recovery have been identified, leading to the definition of preliminary mission requirements at the system and subsystem level. Consequently, the consolidation of the return mission design was possible and optimised trajectories have been defined for the DRL and RTLS scenarios.

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The Intermediate eXperimental Vehicle (IXV) is a suborbital re-entry demonstrator that will be launched in 2014 focusing on the in-flight demonstration of a lifting body system with active control surfaces. The Mission Analysis and Flight Mechanics of such mission is an End to End process from lift-off to splashdown which must be compatible with requirements and constraints in terms of Flying Qualities, aerothermodynamics, layout restrictions and capabilities, structure, launcher performance, Guidance Navigation and Control (GNC), visibility from ground stations and GPS and safety. The results of this design are used to support the specification of the vehicle subsystems, like GNC, Thermal Protection System (TPS) and Ground System. The main products are a database of trajectories, a Centre of Gravity (CoG) box, a trimline and a ground stations visibility network. This paper presents the methodology and results of the validation phase carried out in support to the Critical Design Review (CDR). Validation is a fundamental step not only to verify the suitability of the design but also to obtain the detailed envelope of performances of the mission. Mission and Flight Mechanics validation is mainly based on the performance evaluation against environment and system uncertainties using high fidelity end-to-end models, intensive simulations (Monte Carlo) and worst case analyses. The conclusion is that a feasible Mission Scenario for the IXV Mission from Lift-off to Splashdown has been validated at CDR level, covering Trajectory, Safety, Visibility aspects and a Robust Flight Mechanics solution, which validates the proposed design approach.Nomenclature C mcg = Pitch moment coefficient derivative with respect to the angle of attack at the CoG C n,DYN = Dynamic C n (yaw moment coefficient derivative with respect to the angle of sideslip) d e = elevator deflection d a = aileron deflection d f = individual flap deflection L/D = Lift to drag ratio

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Landing a spacecraft on Mars with high-precision landing accuracy is a challenging task. One of the most critical phases is the atmospheric entry for which an efficient guidance has to be designed in order to achieve a desired landing accuracy better than 10 km (high precision) or better than 1 km (pinpoint landing). In the frame of the Mars Robotic Exploration Preparation (MREP) Programme of the European Space Agency, a study is currently being carried out to perform an end-to-end optimisation and GNC design for high precision landing on Mars. This paper presents the results of the guidance concepts tradeoffs and design phase and the performances obtained by the candidate guidance solution selected. As a further step, an innovative integrated EDLS/GNC sizing tool (ESAT -EDLS/GNC Sizing and Analysis Tool) was designed and implemented to support system trade-offs decisions and optimisation in a multi-disciplinary design environment. The overall system performance obtained confirms that a feasible solution in terms of achieved payload for the required landing accuracy exists and that the entry guidance plays a key role in making the EDL system capable of reaching the target landing site with the desired accuracy. NomenclatureDGB = Disk Gap Band D-E = Drag-Energy EDLS = Entry, Descent and Landing System EIP = Entry Interface Point ESA = European Space Agency ESAT = EDLS/GNC Sizing and Analysis Tool GC = Guidance and Control GNC = Guidance, Navigation and Control L.A. = Landing Accuracy L/D = lift over drag MDO = Multi-disciplinary Design Optimization MOLA = Mars Orbiter Laser Altimeter MPL = Mars Precision Lander MREP = Mars Robotic Exploration Programme