This paper deals with the aerodynamic and aerothermodynamic preliminary design activities for the vertical takeoff hopper concept performed in the frame of the Future Launcher Preparatory Programme of the European Space Agency. The reentry scenario with the corresponding loading environment for the proposed vehicle concept is reported and analyzed. The hypersonic aerodynamic and aerothermodynamic characteristics of the vertical takeoff hopper are investigated by means of several engineering analyses and a limited number of computational fluid dynamics simulations in order to assess the accuracy of the simplified design estimations. The results show that the difference between Eulerian computational fluid dynamics and an engineering-based design is smaller than 10% for aerodynamic coefficients, whereas a margin of about 30% has to be taken into account for what concerns the aerothermodynamic results. The final results applicable for the prosecution of the launcher design activity are that, at the condition of peak heating, the vehicle features a nose stagnation point heat flux of about 500 kW=m 2 and an aerodynamic lift-to-drag ratio of about 1.2.
Nomenclaturealong wing chord running from leading edge, m D = aerodynamic drag, N F = aerodynamic force, N H = altitude, m L = aerodynamic lift, N/fuselage length, m/chord length M = Mach number/aerodynamic moment, Nm Q = integrated heat load, MJ=m 2 _ q = convective heat flux, kW=m 2 q = dynamic pressure, Pa R = radius of curvature, m Re = Reynolds number Re=m = unit Reynolds number, 1=m S = reference area, m 2 T = temperature, K t = time, s v = velocity, m=s x = distance along vehicle forebody running from nose, m = angle of attack, deg = angle of side slip, deg = leading-edge sweep angle, deg = density, kg=m 3 = fuselage meridian angle, deg Subscripts eff = effective N = nose ref = reference sp = stagnation point WLE = wing leading edge w = wall Y = pitching moment 1 = freestream conditions
