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CIRA, the Italian Aerospace Research Centre, in the framework of the national space program has carried out a feasibility study of a future re-entry spacecraft concept with automatic re-entry and landing operational capability. Such vehicle will be injected in a LEO orbit (i.e. 300km) by the VEGA launcher to execute few revolutions around the Earth and then perform an automatic re-entry flight. After de-boosting by the VEGA AVUM upper stage the vehicle will execute an autonomous flight from hypersonic to subsonic regimes allowing terminal area energy maneuvers, approach and landing on conventional runways. Different challenging design and technology performance shall be fulfilled by the vehicle configuration, materials and functional architecture. In particular, the vehicle shall exhibit improved aerodynamic and maneuverability characteristics together with innovative GNC approach allowing more flexibility in the re-entry trajectories, as compared to typical lifting re-entry vehicles.

The present paper is focused on techniques aimed at including control law dynamics into linear flutter equations. An UAV in an unconventional joined wing configuration has been taken as test case for setting up the methodology. For this test case the pitch control law is included in the flutter equations, obtaining a model to perform closed loop flutter calculations. The extra degrees of freedom of control surfaces and additional terms due to the control law have been modeled by means of dynamic sub-structuring approach. The equations governing the dynamics of the control law are added to the aeroelastic stability equations after a suitable manipulation based on a derivative approach. The interesting aspect of the work is that the control law can be modeled as six external matrices to be properly assembled into the aeroelastic system. This is advantageous when already written codes for flutter evaluation are available since the requested modifications are minimal.

The present paper is focused on techniques aimed at including hydraulic servo-actuator dynamics into linear flutter equations. A large aircraft has been taken as test case for setting up the methodology. The hydraulic servo-actuator pertains the elevator. Without losing generality symmetric modal association has been used. The extra degrees of freedom of the control surface and additional terms due to the actuator have been modeled by means of dynamic sub-structuring approach. Starting from a set of flutter analyses performed at different stiffness values of the control surface rotational degree of freedom a minimum requirement for the elevator stiffness is obtained. This value has been used for a preliminary design of the hydraulic servo-actuator, allowing its subsequent dynamic characterization. The next step has been the inclusion of the servo-actuator dynamics into flutter equations (open loop flutter).

In this paper an automatic procedure aimed at preliminary designing an oleo-pneumatic landing gear strut for a light unmanned aircraft is presented. The whole work is motivated by the necessity to design the landing gear of the unmanned aircraft X-MALE, currently in development at the Italian Aerospace Research Center (CIRA SCpA), according to the Airworthiness Requirements NATO STANAG 4671 (USAR). The landing gear preliminary design is usually carried-out by responding to requirements of maximum overall dimensions, maximum weight and maximum attainable load factor. While the oleo-pneumatic mechanism depends essentially on the maximum vertical load factor and on the strut length, thus on the oleo-pneumatic efficiency, the structure is generally sized by combining the design vertical loads with the maximum expected horizontal loads.

In this paper a method to evaluate ground loads due to the impact of a high aspect-ratio wing of an unmanned aircraft at landing is presented. The aircraft for which the loads are calculated - code named HAPD (High Altitude Performance Demonstrator) - is an unmanned aircraft under development at the Italian Aerospace Research Center. It is a scaled performance demonstrator of an 80m-wing span High-Altitude & Long Endurance Unmanned Aircraft, born as a risk mitigation project of its full scale parent. Its peculiar feature is a Joined-Wing configuration. The joined-wing configuration is a good compromise between high lift, low weight and low global flexibility when compared with a conventional configuration with the same aspect-ratio. HAPD configuration consists of a mainly straight Front Wing and a Rear Wing with negative sweep and dihedral angles. The junction location is at about 70% of the FW semi-span, which was chosen as the best location in the preliminary design phase.

Aim of this study is the development of an integrated methodology able to support the structural design of an UAV demonstrator, having a joined-wing configuration and a high structural flexibility. HAPD (High Altitude Performance Demonstrator) is an UAV multi-mission, very light vehicle under development at CIRA, the Italian Aerospace Research Centre. HAPD structure is redundant as regards constraints, so the internal forces depend upon the stiffness distribution. In addition the evaluation of flight loads has to be performed without neglecting the flexibility of the structure. Because of the extreme unconventionality of the configuration the design of the primary structures has been widely affected by aeroelastic analyses.