It was assum ed that the use of a sphe rical b earing at each en d would preven t bendin g moments from   being appli ed to the a ctuator. Aerodyna mic a nalysi s deter mined t hat skin f riction drag parallel to the fla p  surfa ce contributed only n egligible l oads to the flap system. Th e requi red flap driving mo ment transl ated  into force of 44,482 N (10,000 pou nds) at the cho sen EMA attachment point i f the maximum flap force   were to be experie nced. 

The co ncept sele ction  and the kin ematic layout were  frozen befo re the detailed  design of the actuato r or  vehicle  struct ure were fully unde rstood due to the  long lead time requi red for fab rication of the flap s. In  particula r, no  satisfacto ry concept or te chnology was ide ntified f or the  therm al seal where the a ctuato r  penetrated th e fuselag e skin. Accomm odating su ch motion within  the extremely limited volume and while  exposed to the dema nding environ ment led to design problems for the thermal seal that  were n ever  satisfa ctorily resolved fo r this configuration. Also, as de sign of th e linea r actuator m atured, it becam e  appa rent that  insufficie nt clearance exist ed between t he lower p art of the actuat or an d the fu selage lo wer  longe rons, which were already in prod uction. Even afte r less tha n optimal mo dification s to the fuselag e  structure, there wa s no ro om left for growth in the actuato r size or for acco mmodatin g an adequ ate  thermal seal based on ava ilable technol ogy. 

A more deta iled con cept eventually emerged tha t consi sted of a telesco ping tube su rrounding the  actuato r lower rod and m ounted in a ‘ shoe’ that co uld slide ove r a slot cut in the  fuselag e. However, eve n  after exten sive survey s of i ndustry, no  suitable te chnologie s were found fo r implementing  this slidin g and  telescopi ng therm al seal. The p rinciple probl em was finding mate rials capa ble o f surviving th e necessa ry  temperature within the extremely limite d space avail able a nd of a ccommo datin g the comple x mecha nical  motions in th e high tempe rature, corrosi ve environm ent. 

During thi s time, technol ogy develo pment issu es inv olving the  high temp eratu re CSi C bearings of the  flap  arose. Testin g reveale d that the bearin gs exhibited a far highe r than antici pated coeffici ent of friction. Thi s  resulted i n large be nding moments bei ng applie d to th e linea r actuator for which it had n ot been designed.  Reports from MAN Te chnologie in dicate d that the be aring coefficient of friction  could exceed  0.7 or m ore.  Additional te sting l ater fa iled to show any m eaningful relation ship of th e coefficient to  load, speed,  temperature, or oth er environmental and appli cation f actors. Effort s to r educe th e frictio n by modifying the   surfa ce treat ment or fini sh also failed  to gene rate usable imp rovements. M AN Te chnologie informally   reported p romising i ndications of success with a ne w type of prop rietary su rface  treatment, b ut the proje ct  sched ule coul d not accom modate the n ecessary technolo gy deve lopment effort . 

The u nexpe ctedly high f riction of the beari ng m eant that rath er than i solatin g the EMA from be nding  moment s, the  bearing could actu ally ap ply a large m oment at the lower end of the linea r actuator EMA.  Due to the volumetri c constraint s on th e actuato r caused by the layout of th e flaps and aft fuselage   structure, the EMA config uration was forced to a sl ende r design with a very high len gth to diamete r  aspect ratio t hat coul d not sup port be nding load s well. 

The ne wly imposed be nding mom ent could only b e accomm odated with a substa ntial increa se in the  diamete r of the lower lin k rod, thu s drivi ng up the overall size and  weight of the  EMA. It would also have  been  necessary to i ncrease th e powe r of th e EM A to me et re quire ment s under the  additional lo ad.  However, b ecause of the above volu metric constrai nts, there was inad equate room for the necessa ry  expan sion of  the EMA and thermal seals. It thus becam e app arent that the EMA concept  requi red a  fundamental redesign. During thi s time the NASA EMA team began to devise al ternative concepts in  preparation for the antici pated ne cessity of changin g the desig n. 

Existing aft fusela ge pri mary stru cture and  flap design severely boun ded alternativ e EMA layout  concepts. All new de sign concepts also had to ad dress the diffi cult therm al seal d esign probl em. Th e  developm ent time line was also a drivin g issu e, as the EMA redesign effort came ab out ve ry late in the  over-all project schedule. This precl uded sele ction  of de signs or method s that app eared to req uire a   substantial te chnology dev elopme nt effort. 

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