Due to the exposure of the  relatively thin flaps of t he X-38 configu ration to the full re-ent ry environm ent, it  was decided to fabricate them from  a hot stru cture material. A convent ional cold structure flap covered  with thermal prote ction m aterial s was deeme d impractical for the  volumetri c constraints i mposed o n the  flaps. Furth ermore, a s a re sult of the strength limitations of  the sel ected h ot structure materi als, the X-24   driven geom etry of the fl aps, and th e load environ ment expe rienced by the flaps, it was necessa ry to  apply the  actuating fo rce t o the flaps at a point as near a s possible to  the center of p ressure, which  coincided with a lo catio n near th e cente r of the  roughly squa re shaped flap s. This geomet ry dictated th at  the flap actu ator be lo cated  in the aft fuselage in t he region a bove the flap and th at the linka ge conn ectin g  to the flap would p enetrate thro ugh the lowe r skin o f the fusel age to conn ect to the top side of the fla p,  thus expo sing the actuato r linkag e to the re- entry en vironme nt betwee n the flap  and lower ski n. 

The bul k of the X-38 ai rframe co nsists of a cold structure alumi num frame wit h comp osite skin pan els  covered  by a  therm al protection  system (TPS).  The flap s mounted on  the lower side of the aft fu selage  are di rectly expose d to the flow duri ng re-ent ry and consequ ently experie nce p eak su rface tempe rature s  that can re ach as high as 1649 °C (30 00°F). The struct ural layout  of the vehicl e req uired th at the entire   flap fit within a thickn ess of approximately six inches, which would h ave been in sufficient for a   conve ntional cold st ructure  flap of sufficient stiffness and cove red with a TPS of the requisit e thickn ess.  Consequently , the flaps we re con structed of a hot stru cture carb on silicon carb ide (CSi C) ceramic matrix  comp osite  without a TPS. MAN Technologie  of Germany supplied the  flap s via a E uropean Sp ace  Agency / TETRA (TE chnologie s for fu ture sp ace T RAnsp ortation system s) spon sored join t venture wit h  NASA. 

To accomm odate the interface of t he hot structu re flaps to the col d structure airframe, MAN develop ed a  high tem perature jo urnal beari ng mate rial b ased on CSiC, and designed a system of CSiC beams to  support the fl ap at the hin ge line be arings while a ccommodatin g the differential t herm al expan sion ratios  of the hot an d cold structu res. T he hot structure be ams were su rround ed by th ermal b arriers at the poi nts  where they penetrate  the TPS of the cold st ructure. This preve nted imping ement of re -entry gases on the   cold structu re. In ord er to simplify the hi nge lin e inte rface seal s, the flap hi nge line bearings and supports  were lo cated  outsid e the airfram e TPS  such that the se als a re entirely stati c and do not flex during  rotation of the flaps . 

Since the X-38 wa s an u npowered gli der inten ded for long -term  storag e on orbit, it does not have an   auxiliary power sy stem or hydraulic system such as the Space Shuttle. All of the flight c ontrol surfaces are  actuate d with  electrome chanical a ctuators (EMA) th at are d riven by dire ct cu rrent from Nickel-Cadmi um  batteries. Prior development program s (EPAD et al) i ndicated then state of t he art EMA technology could  be su ccessfully applied as a desi gn sol ution for X-38.  Further confi dence was g ained throug h X-38 drop   testing (V132 , V131R) p erformed  at Dryden Flight Re search Ce nter using single strin g, linear output  EMAs. At thi s point, it wa s evid ent tha t the ava ilable pul se-width modul ated, brushless DC m otor and  motor d rive tech nology was a dequ ate to drive th e type of power needed to control th e V201. The   techni cal concerns shifted to the optimum architecture  that should be employed to facilitate human rating.  The h uman rating requi rements drove  the system  to be fully t wo-fault tole rant mea ning that all flig ht  control sy stem requi rements for loa d and rate be met even after two failu res. The mo st fundame ntal  architectu re redun dancy deci sion involv ed the choi ce betwee n an active-stan dby system a nd an activeactive syste m. It was cle ar that an a ctive standby system would be sim pler in terms of softwa re and  control, but would re sult in a more extre me cont rol surface tran sient during a ch annel failu re. The active - active syste m promi sed redu ced failure tra nsients at the expen se of sop histicated  control  software that  would ma rshal four hig hly responsive channels operating sim ultaneou sly while  forced  into lock step by  being  torq ue sum med through  a bull gear. Ultimat ely, the latter path (active-active ) was chosen and  successfully develop ed. 

Design Ev olution 

The o riginal actuator de sign consisted of a long li near ball screw with the top e nd anchored to a spherical  beari ng near the top of the  aft fuselage structure an d attache d at the lower e nd to the high t emperature  spherical be aring n ear th e cente r of the  flap (Figu re 3). Du e to the kine matic l ayout, the linear a ctuato r  would have p enetrate d the  skin at a dyn amically chan ging lo cation,  compli cating the thermal seal de sign.  378