COpy NO . 0:2.- Report MDC E 0056 15 December 1969
A TWO-STAGE FIXED WING I SPACE TRANSPORTATION SYSTEM
FINAL REPORT Volume I Condensed Summary
Contract No. NAS 9·9204 Schedule "
[ [
(COO!) 31 (CATEOOI'O Report MOe E0056 Volume I 15 December 1969
FOREWORD
This report in three volumes, summarizes the results for a McDonnell Douglas Phase A study of a Two Stage-Fixed Wing Space Transportation System for NASA MSC, and is submitted in accordance with NASA Contract NAS9-9204 Schedule II. The three volumes of the report are: Volume I - Condensed Summary; Volume II - Preliminary Design; Volume III - Mass Properties. This is Volume I which presents a condensed summary of our study results. This was a five month study commencing 16 July 1969 with the final report submitted on 15 December 1969. The objectives 'f the study were to provide verification of the feasibility and effectiveness of the MSC in-house studies and provide design improvements, to increase the depth of engineering analyses and to define a development approach. The preliminary design was to be accomplished in accordance with the design requirements specified in the statement of work, and with more detailed requirements provided by MSC at the outset of the study. After the study had progressed to about th, mid-point, NASA redirected the study from a baseline 12,500 1bs payload orbiter to a 25,000 1bs payload orbiter wad changed the payload compartment size from 11 ft diameter by 44 ft, long to a 15 ft diameter by 60 ft long. Directly after this change the program was interrupted so that MDAC could respond to special emphasis requirements imposed by the September Space Shuttle Management Council Meeting. ..I In the interest of clarity and expediting the report, the additional con figurations studied will not be covered in the document. Only the configuration [ having a 25,000 lbs payload in a 15 ft diameter by 60 ft long payload compartment is described in this report. However the information on other configurations had [ been transmitted previously to NASA as the work progressed. The study included eight tasks: Flight Dynamics Analysis, Thermal Protection [ System, Subsystem Analyses, Design; Mass Properties Analysis; Mission Analysis; Design Sensiti~ity Analyses; and Programmatic Analyses. The study was managed and supervised by Winston D. Nold, Study MP-nager of I McDonnell Douglas Astronautics Company - Eastern Division. NASA technical direction was administered through James A. Chamber1in,and contractual direction was provided I by Willie S. Beckham from NASA Manned Spacecraft Center.
I MCDONNELL DOUGLAS ASTRONAUTICS Report MDe E0056 Volume I 15 December 1969
TABLE OF CONTENTS Section Title Page FOREWORD •• · . . · . . i !I 1. Summary ••• · . . . . . · . . 1 2. Design Characteristics · . . . . . · . . . . . 1 1] 2.1 Orbiter • • • • • • • • • · . . · . . . . . 2 2.2 Booster ...... 3 1J 3. Structural Design 3
4. Aerody~amics and Performance • 6 4.1 Aerodynamics • • · ...... 6 4.2 Performance . . . 8 fl 5. Thermal Protection System ...... 10 11 6. Propulsion · . . . · ...... 13 7. Integrated Avionics 16 II 8. T~chno1ogy and Development Plan · ...... 18 IIu
LIST OF ~FFECTIVE PAGES
i chrough 11 1 through 18
u n n u
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MCDONN...... DOUG ... AS ASTRONAUTICS u I Report MOe E0056 Volume I I 15 December 1969 1. Summary - The growth of future an integral tank structure with manned space expoloration is dependent associated frames to pick up the con I upon the development of a reusable space centrated loads. The fan cruise transportation system with operational engines are fixed in the forward practices similar to present day air fuselage which aids in balancing the craft procedures. Such a system could vehicle and simplifies the installa I achieve a dramatic reduction of oper tion. The primary heat protection is ational costs and allow a rapid expan provided by silica cloth faced hardened sion of space flight. insulation and pyrolized carbon I laminate .omposite. A two stage configuration satisfy ing these requirements has been We have concluded that this con I conceived by NASA-MSC. An important cept is a viable configuration. The feature of this configuration is that technical analysis and design results both the orbiter and booster have fixed bear this out. As appropriate, per wings and tail and look similar to tinent analyses and data generated I conventional aircraft. Tre fixed wing by the NASA-MSC in-house study is in provides good subsonic cruise and cluded in theceport. horizontal landing characteristics I which are very similar to present day 2. Design Characteristics - The high performance aircraft. reusable space shuttle described in this report is a tl-/O stage fixed wing system. The ability to enter the Both stages - booster and orbiter - have I atmosphere with a fixed wing is made a similar shape. The booster is powered possible by configuring the vehicJe by ten (10) high pressure bell engines to be aerodynamically stable at high of 400,000 lbs. sea level thrust. The I angles of attack of approximately 60°. orbiter has two (2) similar engines. This effectively exposes only the The system is sized to carry a 25,000 bottom surfac~ to the entry heating, lbs. payload to a 270 NM 55° inclination I which in turn is also considerably orbit. Both vehicles are capable of reduced because of the low planform low level horizontal flight with forward loading. Sufficient analysis has been fuselage mounted turbofan engines. Figure accomplished to show that this concept 1 shows a three view of the configuration I is feasible. A vehicle can be aero and presents some of the pertinent dynamically configured to have a characteristic data. Figures 2 and 3 hypersonic through subsonic velocity present a summary of the weight I high a trim point and also be able to breakdown and mission history weight fly subsonically at a trim low a. for the orbiter and booster. [ A reaction control system is used to provide on-orbit attitude control VEHICLE CHARACTERISTICS and terminal rendezvous and docking tJ. translation V. The RCS also pro PAYLOAD: 25.11. La (UP & 0011' I I vides attitude damping and roll atti PAYLOAD lAY: 1HT II • T tude control for lift vector orienta GROSS LifT ~F IT: :.1154. LI CROSS RAlGE: ZJlI •. 226 FT tion about the velocity vector during =1 MANEUVUIIIG 'V: _ FTISEC I entry. MAX LlO HYPEII10IIIC ~1 ORIITU:U lOOSTER: 1.5 Designs of bo~h stages incorporate LAlDIIG SPEED: \JI UOH conventional structural design tech MAX LID IlIS011C I ORIITER: UD niques. The fixed wing is of conven lDOSTER: 7.15 tional construction, except for the LANOIIIG IE IGHT heat protection. The fuselage uses DRillER: 151 .... LI I IOOSTER m,U'LI Figur. 1
I MCDONNELL DOUGLAS ASTRONAUTICS Re~ort MDe E0056 Volume I 15 December 1969
WEIGHT SUMMARY - 25K PAYLOAD WEIGHT SUMMARY - 25K PAYLOAD
GROUP ORBITER BOOSTER CONFIGURATION DRBITER I BOOSTE R BODY STRUCTURE 39180 92.700 WING 14.100 31.410 LANDED WEIGHT LESS PAYLOAD 133.840 i TAIL 6.140 16.640 25.000 THERMAL PROTECTION 18.450 30.130 PAYL0AD LANDING GEAR & DRAG CHUTE 6.400 12.150 LANDED WEIGHT :58.840 311310 MAIN PROPULSION SYSTEM 16.600 15.885 fLY·HOME PROPELLAN' S 1010 80.000 AIR BREATHING ENGINES & SYSTEM 14100 30510 FLUID LOSSES lUlu 11420 RCS &TANKS 2.500 3.500 ON'ORBIT MANEUVER Itt.V - 2000 FPSI 28.460 AEROOYNAMIC CONTROLS 2.100 4.650 202.280 HYDRAULIC SYSTEM 1.590 2.930 ORBIT INJECTION WEIGHT 400.000 ELECTRICAL POWER SYSTEM 3.280 3.000 INJECTION PROPELLANT IW-IS.!lli5 FPSI G&N. INSTRUMENTATION. COMMUNICATIONS 4.260 2.60~ SEPARATION WEIGHT 602.280 414.130 CREW STATION & CONTROLS. & ECS - - RESIDUALS 1.540 3.400 BOOST PROPELLANTS IW - 14.635 FPSI 1.837.110 RESERVE 600 1.200 :REW & EQUIPMENT 600 0 0 STAGE LlFT·OFF WEIGHT ~O2.280 Z.2~.910 CONTINGENCY ------_.- -~ LANDED WEIGHT - POUNDS 1l3.840 3'.1.310 TOTAL LlFT'OFF WEIGHT - POUNDS 2.854.190 Figure 2 Figure 3
2.1 Orbiter - The orbiter accom Conventional airplane controls are modates a crew of two and has a payload employed for subsonic cruise and landing. capability of 25,000 1bs. to and from A retrac~able tricycle landing gear is orbit. The payload bay is 15 ft. in prOVided. The wing has double slotted diameter and 60 ft. long. The inboard flaps which allows landing. speeds under profile and geometric data is shown in 140 knots on a standard runway. Four Figure 4 and 5. (4) turbofan engines provide the power during this phase of flight.
INBOARD PROFILE - OREirfER
'\ TURBOFAN GIMBALLED ROCKET ENGINES _.\ \ \NVIRONMENTAL CONTROL SYS. - TRACK. TELEM ON ORBIT TANKAGE &COMM.EQUIP. PAYLOAD INTERFACE HATCH \ :\fNEfPAYLOAD ACro"lION MECHAN'" • . PAYLOAD BAY 0~A - . l: ~W-·CjL02
THERMAL PROTECTIVE SYS MAIN PROPELLANT TANKAGE
ReS THRUSTER STA 751 STA 1175 STA 0.0 STA 1410 FWO EQUIPMENT BAY ,'WD INTERSTAGE AFTINTERSTAGE l ATIACH ATTACH Figure 4 2 NlCDONNEL.L. DOUGL.AS ASTRONAUTICS Report MOe E0056 Volume I 15 December 1969
GEOMETRIC DATA - ORBITER main rocket engines to minimize trapped - --I fluid and line losses, (3) the for \I"'.!!.l~.l.l:.. .,""lIlIh~t.! Hothontal rau "ao_try l;tOIl8 W.laht t,O] ,:110 lh. Watted At .. I,))'" ft.' I ~.tllfY We1lht 11' ,<:flO lb. itO) ft2 ward interstage attach point 1s located landing .l"I~ht l~" ,8/i0 lb. [.poal!d Ana 177 ft} Span (b) Ii) ft at the orbiter gross weight CG so that ~.hlch G.v.!!..!..!.l A.pact Rat 10 CAR) 10.68:1 Total J'rojected l. L Sweep 'Q' stage separation is primarily transla Planfon! Ana Taper Ral..t> J~": 1 Root Chgtd (CR) 20. ~ ft tion with a minimum of rotation, and Iod1 (;eOUUy T1p I.:hcrd (Ct) 7,l5 ft Watted }r •• 11,967 ftl M.A.C, (el 1lI.9 It KL Valu... 66.1080 ft 1 Airfoil Sac lion NACA 0012-64 (4) the electrical power equipment, \,UI'th 14f1 ft F.hvator 56% C it !I bolla. Watt"d Area J,O.17 Itt EleVator Dlrfhctlon :!:. 40~ batteries and fuel cells are located in
~n.i.._I.0.!..o~l ~.t.i:.!.!.l!!.!...i!.!.-o~ the forward section to aid in provid Wettd Area l,f)"'l ft.' Wattad Aua 910 ((2 1,850 fll Theuretlcal Area IoH ftl ing a favorable CG. b.po ..d Ana 1,146 ft} Eapu ••" Aua 1,55 fl2 S,.n (b) 111. ~ ft Span (b) 21.2 ft A.pacl l.etl0 (,u) 7:1 A.pact Ratto ("1\) .":1 Dthedul Mile L. E. Swnp l. f .. S_ap ",.' Tapar RaUo '".472: 1 2.2 Booster - The booster is r.per ".t .., .HJ:l Koel Chord (LR) 29.2 fl M.A.C. (i:) 17 • ~ [l Tip Churd (erl lJ.7~ It ~imilar in shape to the orbiter with Root Chord (CR) 24.1 ft M.A.C. (~) n." ft tip Chgrd (t:TJ 8.S It Airfgll SlIl;tiun NACA 0012-6 .. a fixed wing and conventional aero Airfull Section lOt C X b RAGA OO14~6" Rudder Oe,uctlon ! 2~· Ah:!l~O;!c:~~:!y t) dynamic controls for horizontal flight. It tip MACA 0010-64 Flap •• Double Slottad ]0% C X 60% b upoled The booster has the capability for fl.p. Mov ...nt Hall ". Aileron. Ht C X 101 b expoaeg both manned and unmanned operations. Ailllron. Dd hl;tion :t 10· Figures 6 and 7 show an inboard Figure 5 profile and present some of the geo metric data. Two (2) bell nozzle high pressure boost engines using 102 and LH pro The booster is powered during pellants provide p0wer. for initial or ascent by ten (10) bell nozzle bit injection. TIlese same engines 400,000 lbs. thrust rocket engines using L0 and LH2 propellants. Six (6) used in a pres~~re fed mode and draw 2 ing propellant from separate tanks, turbofan cruise engines, installed well provide the power for on-orbit trans forward to aid in CG control are pro lation and deorbit. vided for flying back to the launch site giving an all azimuth l~unch The structural design approach capability. The structural uesign uses the main propellant tanks as an of the booster is similar to the integral part of th~ fuselage struc orbiter but somewhat simpler because ture. The tanks are aluminum and form no payload bay discontinuity is pre ed into a double bubble design with a sent. central web. Titanium longerons, frames and skin make up the rest of The thermal protection consists of the fuselage structure. The wings, only hardened compacted fibers over stabilizer, and verti~al fin use a the high temperature regions of the titanium integral stiffened skin spar fuselage and aerodynamic surfaces. and rib construction. 3. Structural DesiAn - The pri The thermal protection consists of mary structu=al design of the orbiter hardened compacted fibers on the fuse is illustrated in Figure 8. Basic lage bottom and sides, and the under body bending/shear structure is made up side of the wing and stabilizer. Pyro of upper longerons adjacent to the PllY lized carbon laminates are used for the load compartment and the propellant fuselage nose and the leading edges of tank structures below the payload join the wing, stabilizer and vertical fin. ed by fuselage side skin panels. Two integrally stiffened cylindrical tank Other key features shown on the in shells are joint at a common keel web board profile are: (1) the forward in a "double bubble" arrangement. Side location of the turbofan cruise engines panels are single skin, stiffened by to provide a favorable CG, (2) on corruga tions • Thede pane Is and pay·· orbit propellant located aft near the load doors are thE' upper surface of
3 MCDONNELL DOUGLAS ASTRONAUTICS Report MOe EOOS6 Volume I 15 Decem~er 1969
INBOARD PROFILE - BOC~TER ! ! ! YOR/ILS LOCALIZER EQUIPMENT BAY ANTENNA COCKPIT EQUIF '.ENT BAY DME ANTENNA JET ENGINES 'IHF-OMNI ANTENNA ATC ANTENNA RCS ORBITER ATTACH POINTS II
j 11 i
..:- .. ..-- _ . - . -MAt.! PROPELLANT ..... YHF-OMNI TANKAGE RCS ANTENNA ROCKET II LADDER ENGINES DIE ANTENNA ATC ANTENNA Figure 6 II
GEOMETRIC DATA - BOOSTER Frames are titanium to minimize he.at n conductance to the tanks. The upper Yehle!. Wd,hh ~.!!!l...9.OMtr' eron Vdlht l,Bl,tiD lb. W.ttd Aua ],216 tt Z structures are warm during launch and Entry "-tlhl 4110.710 l~. 2 .l~.1 f ,l Laa4i1"1 Wellht )17,)10 lb. b,a•• d Ar •• 1,6011 HZ Span (III' 101.0 It entry, and, also are titanium for good Vehlcl. Ce.. nz A',act a.ul (Al) 4.1:1 Tolal ... oJ.ct. ... L. [. sv..p 10' strength/weight ratio at elevated tem 'lanh", Aua Ta,u laUo .397 :1 loot Chord (C.) 21.6 it peratures. loeb c._try T1p Cho,. (et' 11.310 ft Vett ...... 21,'00 tt Z M.' C. (l) 22.!l ft MI. Valu. 164,)10 ttl .Ur(~i1 Section NACJ. 0012-0" tA·lth 211 ft Uavator )62. C X b The forward fuselage st~Jc,ura1 H Ion_ Watt ... Au. ~.'''O (,2 navalor ,"all."tlo"- ~ 40·
yuucal Tail Caoa.try shell is titantum single skin stiffen Win:.~:-!:!. ),401 ftl Watt,4 Ar .. 2.0'1 .. tt2 naouUul A.... J ,100 tt': 1,041 tt" ed by corrugations and frames, and bp•• " Ar.. 2,704 ~t2 Expo ... Ar .. 1,0107 H2 Span (~) ao ft Span (b) )0 ft forms the M.L. except where non A.,act lAUo (Al) ',92:1 bp.~,;t hUo (AJt) ,9:1 n Dihedral bile 7" L. E. Sw .., 1,,). structural surfaces exist, such as L. r:. Sweep 1/0- laper btio .1I6J:l Ta,er bUo ,H)-I Aoot Ctu~r4 (ra ) lo).l It M.A.C. (l) 21".111 ft Tip Chord (el) 20 It engine and nose landing gear doors. loot Chord (ea> 1'" ft H ..... C. (c) 12.7 ft Ti, Chou (C1) 12 ft Airfoil Section IUCA 0012-"" Intercosta1s and frames are transi Airfoil 'ut1~ av44u ;01 C X h tion structures between the forward .Airfoil Section It Ti, MACA 0010-_'" fuselage and the propellant tank. ___nap·_·_h_.'_(b_Od_,_t_)•• DOIoIHe UoUe4 __ )Ot C 1 bot__'_'-J b npoiled___ __. ____ nap. MoY_nt Mu ~_C_'A_OO_14~!I. '_~_d'_'_O'_fl_«_"_U" ~'J AUerOft. 2~1 C X )01 b U(H).~ . AUuOfta Deflection :!: 2;)- 1 The hardened compacted fiber is u L bonded directly to the forward fuse- Figur.7 lage shell surface aft to the propel lant tanks. In the main area twenty I the fuselage. Tank shell structure inch long HCF shingle panels form the is aluminum for compatibility with bottom and the sides up to approxi propellants and protected by moldline mately six feet above the chine lines. Theme,l Protection System (TPS) shing Single thickness beaded titanium panels I les. Shell stiffening frames spaced form the surface between the HCF shingles at 20 inc~ intervals also support the and fuselage structural side skins. TPS, upper side panels and longerons. HCF is bonded to fiberglass honeycomb I
4 I ItIICDONN.LL DOUGLAS ASTRONAUT. Report MOe E0056 Volume I 15 De;ember 1969 panels which distribute surface pn~s supporting cabin pressurization and nose sure loads to small lateral shingle gear loads. support beams. The beams are attached to the tank shell stiffening frames by The orbiter wing is attached to titanium links spaced at approximately the fuselage at three major frames in 24 inches across the fURelag~. Re the plane of wing spars at stations moveable Pi shaped elements attached 391, 972 and 1024, and to the keel web to the beams retain the shingles and in the plane of the wing t rib. Normal ~rovide a gap for thermal expansion. wing loads and symmetrical wing torque are supported at the frames and drag The two orbit~r boost engines are loads are sup~orted at the keel web. supported by a tripod arrangement of linkage thrust structures for each engine. The primary two cell wing box is Linkage loads are transferred to the made of 6Al-4V titanlum wi th 1n'·.egrally keel web, upper longerons and frames at stiffened skins of conventional arrange stations 1635 and 1717. The frames also ment, Figure 9. The main box is pro serve as main support elements for tected from reentry hedtin~ by external vartical and horizontal tails. insulation (HeF) bonded to the lOllier surface. The thickness of the HeF Ls Turbofan engines are supported on established to not exceed a bond line longitudinp.l intercostals attached to the temperature of 500°F. The upper wing forward fuselage shell and by bulkheads surface experiences temperatures of at stations 320, 362 and 400. The ~ess than 800°F, and, therefore, is bu1khec.ds also serve as primary structures not insulated. The orbiter wing lead ing edge is constructed of carbonI STRUCTURAL ARRANGEMENT - ORBITER I ! I j TITANIUM [ONGERO~ DOOR-,." __ ~ '\ ,t /,' I ~TITANIUM STRUCTURE 1I \ .1- \ I BEADED TITANIUM SHIN~I_E ; ... . rfl/HCF L-_.... -' -u '~,.111 .: \ SHINGLC.S A-A B-B---
TPS SUPPORT UPPER LONGERON~t: r B STRUTS-v ID ~tu~~'~I~~~~~~~~~~~~~~~LU~~~a~~ , I
L.:A TPS SUPPORT C BOOST PROPElLANT STRUCTURE TANK (ALUMINUM)· AFTINTERSTAGE FWD INTfRST AGE CONNECTOR CONNECTOR- Figur.8
5 MCDONNELL DOUGLAS ASTRONAUTICS Report MOe E0056 Volume I 15 December 1969 carbon composite honeycomb sandwich REPLACEP,BLE "SLIPPER" LEADING EDGE CONSTRUC1.ON material that serves as structure and Sli~ -.er Designed for 10 Flights - requires no additional TPS. The lower side of the wing leading edge in the inboard region is the area of the most severe heating. In this area the carboni , ",_ - ' ',_ I carbon composite will exhibit an oxida '" ,,-:------• " -" < "'8 tion erosion with each entry. To keep "'. " - from having an excessive thickness and C' """? - - ,'~O~rl ..... • I~'JV rAR80N (U8.JN ending up with a distorted airfoil, a HoM~C,lIIIfI replaceable slipper leading edge is H('"~ propose1 as shown in Figure 10. The :1 slipper thickness is sized to allow : : ItP"I'rt"fHf II SUPP[R about 10 entries, and then can be 8~' 1 easily replaced. The titanium struc tural box is insulated from L.E. radi e • ative heat by a layer of HCF on the front spar. 4. Aerodynamics anrl Performance The booster fuselage is similar in concept to the orbiter fuselage. The 4.1 Aerodynamics - AerodynamIc main propellant tanks are "integral" analyses have been performed for each a1uminu~ body structure and carry over of the var~ous flight regimes from all vehicle loads, as well as, internal liftoff to landing for th~ nominal pressures. The forward fuselage primary mission. Miseion considerations have structure is the outer shell which con resulted in the establi~hment of sists of stiffened titani.m skins an, several aerodynamic configuration frames, pt'otected from ascent and re requirements: (1) high angle of attack entry heating with externc HCF similar (ex - 60°) trim capability thl'oughout the to the arrangement on the orbiter for hypersonic/supersonic portio~ of entry ward fuselage. The booster wing leading with controls fixed; (2) static stability edge experiences lower temperatures, in pitch and yaw with neutral stability relatively, and is a titanium struc in roll; (3) the capability to trim tur~ with extel~al insulation (HCF). subsonically at both high (60°) and low (S°) angles of attack with adequate transitional control; (4) handling WING STRUCTURAL ARRANGEMENT - ORBITER qualities for subsonic cruise, approach and landing typical of present high t --- .~ ;~. 5rA.JI i -- performanLe aircraft. These require ments, in turn, have led to configuration ~f AI" ,dhl. llf II "'C5 ror selection guidelines which C3n bp summarized for entry as: (1) the lo~er '. surfaces of the body-wing-tai1 com
1 bination should be smooth and continuous H L-______r=:- ',;\I:;,:.. ,t '':Ot .~. ~ A. to minimize flow interact",n; (2) pitch trim will be obtain~d ~y camb, :', :\g the flat fuselage bottom fore n and a~_ of the center of gravity in combination with the horizonthl tail; (3) lateral stability will be obtained by wing dihedral (7°); n fJ
6 MCDONNELL DOUGLAS ASTRONAUTICS 11 Report MOe E0056 Volume I 15 December 1969
(4) dir~~tional stability obtained by The Hypersonic Arbitrary-Body differential fuselage side wall angles Aerodynamic Co.nputer Program was utilized for and aft of the center of gravity to predict the hypersonic aerodynamics (i.e., 5° cant angle on the forward for the orbiter and booster configurations. section and straint sidewa~ls aft such The results of this analysis of the that at small angles of si<:aslip flow orbiter as presented in Figure 12, show impingement will produce stabilizing that the orbiter can be trimmed in the moments); (5) reaction control system region of t max (500 to 60° angle of for stability augmentation; (6) low attack) with a center-of-gravity W/SC (~ 50 psf). Similar guidelines (e.g.) location between 53% to 59% of were Lestablished for the subsonic cruise, the fuselage length. The forward e.g. approach and land portion of the flight: limit is the point at which the vehicle (1) fixed wing design with low sweep would trim without an elevator, whereas (14° leading edge), high aspect ratio the aft limit is a stability boundary (AR • 7) with convent~onal ailerons beyond which no stable trim point exists. and double-slotted flaps for landing; For a down el~vator (positive deflection) (2) conventional vertical/horizontal of approximately 25° there are no stable tail, rudder/elevator; (3) sufficient trim points. Also shown is the trim turbof~n power and LID for typical lift coefficient (C ) and lift-to-drag L airplane handling qualities during ratio (L/D). The respec~ive maximum approach a~d landing. Consideration of values are 1.85 and 1.6 at angles of 0 0 these requirements/guidelines, and attack of 52 and 20 • At the proposed various parametric studies have led entry angle of attack of 60°, CL • 1.8 to the selected configuration. Prime and L/D • .5. emphasis has been placed on the orbiter, however, similarity between the orbiter An estimation of the hypersonic and booster results in most of the static and dynamic derivatives was aerodynamic characteristics being cornmon. also made. The data indicate the vehicle is dynamically stable in yaw, pitch, ~ The launch configuration results and roll; however, the vehicle is prp.sented in Figure 11 have been statically unstable in yaw for angles developed using the low speed wind of attack less than 55 degrees. tunnel data obtained from Langley Research Center. These data must be considered as preliminary estimates. ORBITER HYPERSONIC AERO CHARACTERISTICS The drag is adequate for preliminary lot •• "JIG . IUD fll launch trajectory calculations. In addition, t~a negative Cm and positive a Cms tndi~dte an irherent stability exists. LAUl'fCH CONFIGURATION AERODYNA~IC CHARACTERISTICS ~ ::1.. __ , __ -, __ . lot" ". ~_. ~7 1111',1 .. ! "otG~ ___ ~ ____ ~:l ~- k>-'~,,", -r "---T- ~-:'"liclfO"U- ~ .~--~~~~~ -LlnnHCC --"",,"GCG --j- I",.G - 'ulli,.lICa HAllOM - III Figur.12 -T- r 1i -----, . --r- - t ---i- ---1 Subsonic aerodynamic characteristics for the orbiter configuration have been _-Qltlej--- ...------+- derived from a NASA Langley wind tunnel test. These test data were modified to reflect small configuration variations : I "Ct:~' Figur. 11 7 MCDONNaLLDOUGLASASTRONAUncS Report MOe E0056 Volume I 15 December 1969
including nose fineness ratio, tail port aircraft because of the highel size, and horizontal tail aspect ratio drag associated with the base area changes. Modifications were also made and fuselage wetted area. The design to the basic data in the angle of 30% chord double-slotted flaps covering attack range between 45° and 75° to 60% of the exposed span yield ~anding . , account fer the difference between the speeds (1.1 VI) less than 140 kts - ! subcritical test conditions and the and produce gSSa ~orizontal take-off - , super-critical flight Reynoids numbers. characteristics, high C and moderately The resulting orbiter stability is high LID. Figure 14 alsoL shows the shown in Figure 13 for two center of estimated flap effects for landing ~ I gravity locations and including effects (6 - 55°) and take-off (6 - 20°). of elevator deflection. Two separate Th~ techniques used in obtIining these anglp. of attack regions exist for stable estimations yield good agreement with trim (-em). Reentry attitudes lia in a DC-S-6l flight test data and DC-lO wind the high angle of attack trim region tunnel data. anJ adequate elevator control power exists to break this trim point and to Due to the similarity of the orbiter JJ perform the subsonic transition to and booster, the booster trim lift coef the low angle of attack trim region for ficients are nearly identical to those a center of gravity position between of the orbiter. However, the large ~ I 52% and 57% of body length. base area of the booster results in a cruise configuration (LID) of 7.~ ORBITER SUBSONIC TRANSITION compared to S.l for the orW!fer. AERODYNAMICS LRC I.T. DATA ORBITER TRIM AERODYNAMICS 't. r!' SREF • SwiNG SUiSONIC fLIGHT REYNOLD S ,.u.UI lREF • f lUOFL."'DHLlCTIClJlIAUDONl"Clt DATA ADD!T_, DlfllCTlOIIS liT_TID '.011 DC-I-lI' ~C-IO DAT.
CENTER OF GRAVITY - 1]- "., O"LIC1IOII ~ ~~~p..,.:+, OF FUSELAGE LENGTH 01. - DIGI ~d~l'------j ! •. t -+~+-"SO o '---...J.-----"'_.J--~-J :!S~ E~~4°EE 20 ... ~ ..JU ...... J ... 0 II ~so SS 60 XC.G. - , OF FUSELAGE LENGTH
ELEVATOR DEFLECTION :~.-+0 -""f-~I --~Il ----r. T~ 41- -. REQUIREO ELIMINATE lNGL( Of AflACa 0' ... n HIGH. TRIM Figure 14 Figure 13 4.2 Performance - A point-mass It The estimated orbiter low angle launch optimiz~:ion computer program of attack trim 11ft coefficient and was used to compute an ascent trajectory H lift to drag ratio for the subsonic to a 55° inclination, 51-100 NM orbit. flight Reynolds number is shown in The significant parameters are presented Figure 14. The cruise configuration in Figure 15. Note that maximum dynamic IJ data (flap deflection, SaO) is based pressure is about 500 PSF, and that the on wind tunnel data previously discussed. engines were throttled to keep from The maximum lift coefficient is some exceeding 2.5 g's during 1st stage, and II what less than modern airliners primarily 3 g's during 2nd stage. The earth because the standard NACA symmetrical reference {nsertion velocity was 24,965 airfoil used on th~ orbiter and booster ft. per second. The velocity losses (selected to alleviate transonic load due to gravity, drag, back pressure and f I ing during ascent) does not exhibit a maneuvering totaled 5527 ft. per second high CT-max • The maximum lift to drag and giving a nominal ideal velocity ratio is also less than s typical trans- requirement of 30,492 ft. per second.
8 NlCDr'~N.LL DOUGLAS A .....ONAUTIC. Report MOe E0056 Volume I 15 December 1969
NOMINAL ASCENT TRAJECTORY attack at full lift attitude (zero bank h•• 100 NA.III .. '" - SI NA.III. angle). At approximately 260,000 feet, "SS' .---.--,---,--, ~ , the orbiter begins to pull-out due to :;: I the increased aerodynamic lift. The ~ ...a: f\ I :::> orbiter is then bank modulated at ... 1:i co a: constant angle of attack to maintain a ~ __-+-...,.o<:i~-+--+--I A I ' u ~ constant altitude until reaching the c ! ~ V 1\ co velocity of a full lift equilibrium guide entry trajectory. The glide ~ e: i/ entry is then flown until reaching an E~-+-~-~~~ v- I / altitude of approximately 50,000 feet ....> ~ (M = .4) when angle of attack transition is initiated. ~> Q TIllE - SECONDS TIllE - SECOIIlS The advantage of a 60 degree angle Figure 15 of attack is twofold. First, it is near maximum lift coefficient and Computer simulations of several therefore yields low entry load factors possible booster flyback trajectories (1.5 g's) and a low maximum heating 2 have been performed to estimate the rate (62 BUT/ft -sec). Secondly, it structural loading and flyback range is a high drag configuration resulting requirements. The selected control in relative short entry time and low technique involves a negative lift total heat. (180 0 bank angle) during the early low dynamic pressure region to minimize The lateral range capability downrange travel. This is followed associated with 60 degree angle of by a full lift phase to reduce the attack (L/D = .53) is approximately maximum normal load factor. Following 230 nautical miles. In combination peak load factor, the booster is then with the velocity increment available banked 80 0 to turn the velocity vector for return phasing (55 ft/sec), 230 towards the launch site. A typical nautical miles is sufficient for once trajectory is shown in Figure 16. a day return capability.
BOOSTER ENTRY TRAJECTORY ORBITER ENTRY TRAJECTORY Entry from a 270 Na.Mi. Circular Orbit STAGING CINlITIONS ALTITUDE • 2IO.OGO FT VELOCITY • 10.600 FT SEC I·~I 4OOr-_,--_,-_,-..:.F-=.lIG:c;HT. PATH' 6.0 OEG ENTRY CONDITlON AT .00.000 FT .. V1-lS950F'T SEC c 'YI __ 160 5 i :: 3aor__ """'-..:+--+----l t---+--+---+ :: 2Ot----t----F-t---1 • TitANSITlc. •~ L----'_~_-'
",MAX· 4.1 1'5 ", MAX' 0.3 p+--+--+--o-....,..,,,*"---+-.:..:;...;;;;=;'--I ~ ~ £-JO~IIII I :; 4OHY~\-t----1 • TRANSITION , e~ -1 0 I'·. -+------0.:'. :; I;: °IOO~--:200~--,300~-~400~~lOO ~o-~-~----:'':-----:'16 ~ : , : -1.0 RANGE FR i : I o ' Figure 16 12 ~ 4 I U TIllE - 100 SEC Orbiter entry trajectories were Figure 17 computed from a 55 0 inclination, 270 NM circular orbit. A nominal entry is The transition maneuver to bring t:le shown in Figure 17. The orbiter enters angle of attack fro~ 60 0 during entry the atmosphere at 60 degrees angle of to a cruise attitude of about 50 is 9 MCDONNELL DOUGLAS ASTRONAUTICS Report MOe E0056 Volume I 15 December 1969 shown in Figure lB. After this maneuver the cryogenic tank rings with titanium the vehicle flys as a conventional air structural links. These titanium craft can be flown to a horizontal land links are designed to minimize the heat ing. short between th~ exterior panel and the cryogenic tank rings. A low density fibrous insulation blanket made ANGLE OF ATTACK TRANSITION TRAJECTORY of TG 15000 is supported across the tops of the cryogenic tank rings to form a prelaunch purge space between the tank wall and the insulation 11 blanket. Roles in the tank rings permit the purge gas flow to pass from one ring section to the next. On the inside of the hydrogen tank a polyure foam is bonded to the tank wall • -10 -~"£L[VATOft . Q 0 ~ -<-'II i 0.5 - ... 10 i::i' ., -- +--1 S The approach selected for areas ~~ z'loo~ El£VATOR on the wing leading edge where the ~ I I , M .. , 10 :!'D 30 to 511 temperatures exceed 2500°F is a replace 11 TWlE" FRO_ TR~S1TlON" INlTlAnON - SEC " TfII[ fR(,.t TIWISlnott INITlATIOH - SEC able carbon slipper concept. See Figure 18 Figure 10. Inhibited carbon will 5. Thermal Protection System - A oxidize where the temperatures exceed description of the orbiter TPS for entry 2500°F. After several entry flights H at 60° is illustrated in Figure 19. this oxidation may change the aero Pyrolized carbon laminate is used on the dynamic characteristics of the wing. nose cap and wing leading edge regions The replaceable slipper leading edge where temperatures exceed 2500°F. The construction permits a relatively majority of the upper fuselage surface, inexpensiv~ part to b~ designed that upper tail, and upper wing areas are can be replaced when n~cessary. Jj protected with titanium skin because Behind the inhibited carb'n slipper is the temperatures are below BOO°F. a carbon/carbon honeycomb s~ructure Hardened compacted fiber (HeF) insulation in the leading edge that is good for made of silica and bonded to honeycomb 100 flights provided the surface of sandwich panels is used to protect the the carbon/carbon never exceeds 2500°F. lower fuselage area. On the lower wing The slipper consists of a carbon/carbon and tail areas, and on the forward external surface approximately 3/10 of regions of the fuselage, ReF is bonded an inch thick that is backed by zirconia directly to the titanium skin. insulation and attached at local spots to the honeycomb sandwich. These On the bottom and lower side regions attachment points are insulated with of the fuselage a silica ReF material zirconia plugs. The slipper is is used with a 15 pcf density. This cJnsidered only in those areas where ReF has a silica cloth facing that is temperatures above ~500°F are expected. u used to provide increased resistance The actual life prediction for the to rain erosion and servicing damage. carbon/carbon slipper leading edge This facing has a high emittance coating using the worst-on-worst assumptions n of cobalt oxide. The outer layer of for the current heating prediction is ReF is bonded with a film adhesive to estimated at 4 flights. If more a fiberglass honeycomb sandwich. realistic assumptions are selected in Adhesive temperatures are limited to the region of interference heating on IJ 500°F in this design to obtain the the wing leading edge, the slipper maximum reuse capabil~.ty. The honey design thickness is good for roughly comb sandwich panels are attached to 10 to 30 flights. o 10 MCDONNELL DOUGLAS ASTRONAUTICS Report MDe E0056 Volume I 15 December 1969 ORBITER TPS DESCRIPTION HCF material is bonded directly to (a = 600 Entry Trajectory) wing structure, and the bond line temp erature is limited to 500°F. The total ENGINE ACCESS DOORS MATERIAL CODE TPS weight for the orbiter is 18,450 lbs. _ PYROLIZED CARBON This total weight includes HCF, honey LAMINATE; OXIDATION INHIBITED (OVER ZSOOoF) comb panels, structural supports, insul m:l HCF - INSULATION ation blankets, base heat protection, (UP TO 2500oF) and cryogenic foam in the hydrogen tank. HARDENED COMPACTED FIBROUS INSULATION The reference fuselage area and wing BONDED TO HONEYCOMB area protected by TPS are indicated. SANOWICH !ii:i2i HCF - INSULATION BONDED TO TiTANIUII Figure 21 illustrates the booster SKIN baseline TPS. The majority of the c:::J TITANIUM SKIN (UP TO aoooF) area is below 800°F and is protected by OVER INSULATION titanium skin over insulation blankets. BLANKETS Those areas on the lower wing horizontal Figure 19 tail, and the forward areas of the fuse lage that exceed 800°F are protected by Figure 20 summarizes the total the hardened com?acted fiber insulation. orbiter TPS weight distribution along The total TPS weight for the booster is the fuselage, and the chord-wise weight estimated at 30,130 lbs. This weight distribution on the wing. On the includes titanium shingles, HCF, insul bottom centerline, the TPS weight drops ation blankets, cryogenic foam inside sharply on the front 20% of the fuse the hydrogen tank, and base heat pro lage length because the HCF is bonded tection. (Where HCF is bonded directly to the titanium skin rather than to titanium that serves as structural applied to honeycomb panels. On the skin the titanium is not incluc~d in wings the TPS weight is slightly the TPS weight.) heavier at the wing tip (100% of exposed span), because the chord length BOOSTER TPS DESCRIPTION and the leading edge radius are slightly 51 Na, Mi. Insertion 0 smaller than at 50% span. The dash line a = 60 Entry TOTAL TPS WEIGHT. 30,lll LB' indicates the heavier TPS weight in the inboard region where interference heat ing is experienced. In all cases the MATERIAL CODE Q1 HCF - INSULATION (UP TO 2500°F) BONDED TO ORBITER TPS UNIT WEIGHTS TITANIUM CTITANIUM (a = 600 Trajectory) (UP TO BOOoF) OVER INSULATION BLANKETS , INCLUDES OUTER 5 FT • TITANIIJI SHING LES • HCFOVER TITANIUM • INSULATION BLANKETS • HZ CRYO fOM • BASE HEAT PROTECTION Figure 21 Figure 22 presents the orbiter heating rate distribution along the 5~ISPANT fuselage bottom and chine region, and ~ ~!p'f,~i I HCF BONDED DIRECTLY - _ TOP (INSULATION TO II1NG STRUCTURE the distribution around the circumference lJIDER TITANIUM) of the fuselage. These distributions 00 20 40 60 aD 100 0 , FUSELAGE LENGTH are for the baseline trajectory (a = 60°) normalized to a fuselage length of 150 ft. 11 ItIICDONNELL DOUGLAS ASTRONAUTICS Report MOe E0056 Volume I 15 December 1969 FUSELAGE HEATING DISTRIBUTION flow field around the wing. Region (Angle of Attack, a = 60° Fuselage Length = 150 Ft) one moves inboard toward the fuselage as the angle of attack is increased. qREF • 61 BTU/H2·SEC MAXIMUM 2.0 At an angle of attack of 60 0 the SPHERE STAGNATION POINT 0-ill NOSE RADIUS. 1 H 1.0 ...-'-t-"T""I-- fE outer edge of the interference region ... 0.30 LAMINAR HEATING i= I--Il-+-'I- ~ l .... ;:i o.SO is approximately 35% of the exposed .# .';0> :. ®' 0 wing span length. Interference region ~ 1= 0.20 1---t-~-----1 two is caused by boundary layer flow o along the fuselage intersecting with .... ~ 0.101---t---+#---I E the wing. The lower figures show the i l ... e z ~ 0.05 t--t-i-i"-.;;;;:;:;;;;oooi heating rate increase (or the heating fi: 0.10H~"-d----j .... -' rate multiplier) that is used as a :>: ~ BOTTOM l TOP l -' . i, .. 3 0.021---+-++---1 function of chord length for region ~ -' ~ one and region two at three angles of 50 100 0.01 0 1.0 2.0 attack, 15°, 45°, and 60°. Currently , FUSELAGE LENGTH WETTED DISTANCE/BODY WIDTH there is uncertainty regarding extrapol Figure 22 ation of the interference multiplier for the first 15% of chord. However, The heating distribution on the this is the leadit.g edge region of the wing for the design entry condition is wing, where the carbon/carbon replace shown in Figure 23. The data shown are able slipper is used which has been from tests conducted by NASA-MSC on a sized to endure more than one flight. 100% and 40% chord models. The recommended design curve accounting for the flow irregularities and instrumentation difficulties is also WING - FUSELAGE INTERFERENCE HEATING indicated. WING HEATING DISTRIBUTION a~ 60° 5 LRC HYPERSONIC AEROOY NAMIC VI.T. M P,(PSIA) MDDE L SCALE .... 2.5 u 010.35 648 40\ CH ORO 1/7.5 REGION 1 REGION 2 0 •• 600 ffi I o 0 • 60° I ~ (;. 10.3 140 100\ CH ORO 1130 "1" ...... 45° ..... 45° I 1 • EFFECTIVE RADIUS CDRR ECTION ""IE .... 0 15 -n - RECGMMENDED DESIGN CU RVE l- 0 •• 15° 0 • °1 S ~~ E ""zO. • WING TUNNEL CUT OF CA LlBRATION 2.0 "?" 0 ~ ~HORD LENGTH· ISO INC HES ...... ~ 0.2 r~ ::> ....c ~ ~ ~ ..."" o. 1 ....~ z , "" fi: 0.05 ....1!! 1.5 .... ~~ " l- :>: e -' ~ g 0.02 !Ii!"" -' "z i= 0.0 1 r-... .J e .U' WINDWARD .... LEEWARD "~ :>: 1.0 .005 (LOWER SIDE) 0 0.5 1.0 0 0.5 1.0 100 80 60 4 60 ~o 100 , CHORD PERCENT CHORD Figure 23 Figure 24 I} A summary of the data obtained by MSC on wing interference heating is summarized in Figure 24. Two regions The design trajectories from which are indicated. Region one has two the local temperature distribution u zones and it is thought that this shape and the thermal protection system is caused by the fuselage bow shock weights were determined is shown wave combining with the shock wave and in Figure 25. I 12 MCDONNEL.L. DOUGL.AS ASTRONAUTICS I Report MOe E0056 Volume I 15 DeCl~mber 1969 DESIGN HEATING RATE HISTORIES nozzle type engines were selected for Orbiter Entry the booster, and two (2) for the orbit 0 a = -I -, 60 er. All boost engines are throttleable. ~70 ~ .IlMAX· 67 B1U/fT 2'SEC The engines are designed for 100 mission (.) ~ Ascent • QT • 25.097 ClUfF 12 60 1\ life with a 10 hour life between over ~ .1 Fl SPHERE STAG .PT I- 51 Na. Mi. Insertion I ... DETRA. KEMP & RI DOELL haul. ~ 20 r----r-----,,---r---,--, ~50 \ l l II>, II>, ..., 16~~~~~~~~ ..., 40 Figure 27 shows the boost engine l - ~ e feed system geometry. Five 14" dia. '"~ 12~~--~~~~- '"!i 30 ;::: ;::: lines run from the oxidizer tank with e ~ \ :: ~ 20 each line splitting into two 10" dia. Iz <; J \ lines. The line division is positioned ... \ such that a vapor bubble generated by ~ V ;::: 0 .,...... ~~L.,-~---,.!.,...:...... J "- 200 4DO 600 800 1000 an engine shut down will not be ingested ! TIME FROM 400K FT - SECONDS by another engine. Engine isolation ~ valves are located immediately down Figure 25 stream of the line division. The ten resulting lines are then routed to each 6. Propulsion - The propulsion boost engine as shown. ~iffusers are systems required on the booster and used to transition smoothly from the orbiter mission include: (1) a boost 10" dia. lines to the required 14" prop 'lsion, (2) attitude control, and dia. engine supply. Pressure/volume (3) cruise propulsion for both the compensators and gimbal bellows booster and orbiter; and an orbit assemblies are used immediately up maneuvering system for the orbiter. stream of the engines. The oxidizel tank incorporates anti-vortex and slosh The boost engines\were sized baffles. accounting for the 6V iosses during boost, engine out capability, base area contribution to performance, and BOOST ENGINE FEED SYSTEM commonality of engines between the . lfIITHIQIITUaAFflB PRESSURE VOLUIII( eOMPE"SlTM , DIffuSER booster and orbiter. The characteris GI.Al "ELLOIS tics selected are summarized as Figure 26. Ten (1) high chamber pressure bell BOOST ENGINE CHARACTERISTICS BOOSTER OP.BITER l TYPE HIGH Pc BELL HIGH Pc bell LKl fill MIXTURE RATIO 6:1 6:1 a ORAl,. I AREA RATIO 42.S:1 100:1 J (FIXED) '-----, -'--- l02FILL10RAlfil (RETRACTABLE) L02 'HO DuCTS - WEIGHT 4150 4«10 Figure 27 -1, WEI GHTED 3.S The hydrogen feed system is gener IMPUL \E PE NAL TY - SEC 1.6 ally similar, except that due to the NOMINAL THRUST - LB 400.000 463,000 relative close coupling of the hydrogen (S.U (VAe) I tank and the engines, the hydrogen lines GIMBAL REQUIREMENTS are initially fed from a compartmented PITCH ,5° :5° sump. Engine shut-off valves are YAW :il ,1° located at the sump outlets. The hydro ROLL ,1° +1° gen tank also incorporates a multi I cruciform anti-vortex baffle assembly Figure 26 13 MCDONNELL DOUGLAS ASTRONAUTICS I Report MOe E0056 Volume I 15 December 1969 and slosh baffles. The compartmented NOMINAL 1\ V BUDGET sump and the anti-vortex tank baffle are configured so that any vapor bubble generated by an engine shut-down can OIlS RCS not be ingested by another engine. NOMINAL MISSION FUNCTIONS Single point fill/drain vehicle/AGE ORBIT TRANSFER AND CIRCULARIZATION 660 FT SEC interfaces are used for each propellant. TERMINAL RENDEZVOUS 60 FT SEC Initial helium engine requirements are DOCKING AND STATIONKEEPING )0 FT SEC DEORBIT (!MCl 10". RESERVE, 535 FT SEC ground supplied. Upon engine star.t-up, bleed GH and bleed GOX are used to DISPERSIONS 2 PRECEEDING TERMINAL RENDEZVOUS IlO FT SEC pressurize the hydrogen and oxygen tanks DURING TERMINAL RENDEZVOUS 90 FT SEC respectively. The boost feed system GRDUND TRACK AD.MTIIF~TS 55 FT SEC for the orbiter is similar and is VELOCITY IICREIIENT REOUIRED Ut5 FT SEC us FT SEC schematically shown by Figure 28. VELOCITY MARGIN 450 FT SEC TOTAL VELOCITY PROVIDED lOGO FT SEC Figure 29 ORBITER BOOST ENGINE FEED ~YSTEM Res REQUIREMENTS l '~O'[llA'T "0 .... DEG SEC TOTAL. IlIIPUI.Sf. THRUST IT ~[OUIfI[D FUNCTION FT SEC ROll (li-SEel ! PITCH ,.. UI !II / .. ,/... /./..... r ...... : ..... , ORlllTER ! TERttfIHAl RENDEZVOUS 10 0011 - 0 - 10UOO , 1lI0 !fORE AfTl 0001 (OTHER) DOCI\ING 10 0\ I 01 01 IOJ.OOO III ORBIT Aes - 0 0 - 0 1!6000 1010 ROll OIITURBANCE 0 -- ~ 0 ,4, - III ,, "J (BOOST ENGINE OUl) DI\PERlIONI 161 0 - 0 1115,000 IllO !!:[·ENTRY 0 \MAll 10 III 1111.000 1\60.-.-- !rTl)'jIa~-r'-.l BOOST[~ S£"UATrOH !Zl I R(·ENTRY Il) I 77'* IIGIl : I (I BASED ON I, ., C JlOltc , - \ \. J '" :, (,}cc, __ ~_~ :~' 11) ItEQUlftflllENTS NOT DEfIN£D "- ORllTER lItIltANG£MUT IILl BE u!ED PENDING FU_THU DEFINITION : _. - - - _. <' ... --"'-""-"'~::J- .'~ , .~]~ PftOPELLANT ORA. fROM 0111 IT ItCS IUOGE T , (e, ~OUIV"U.NT TO 60 otII fT·L I ROll TORQUE AT 1 1 OEG VAl GI.Al , , Figure 30 , LHz , 1.... ______._ ------The large orbital maneuvers may be satisfied by using one or both of Figure 28 the orbiter boost engines at reducea thrust level, or by adding an addition al engine system, e.g. two additional RL-10 engines. A trade study comparing On-orbit maneuvering and attitude the weight of possible alternatives control requirements are dictated by was made. The lightest maneuver system the nominal 6V budget (~~gure 29) and is obtained with either the use of an the required translational and angular advanced design high Pc bell nozzle acceleration response characteristics engine operating in a pressure fed mode (Figure 30). The initial circulari at 1% thrust, or the use of two addition zation, orbit transfer and retro are al RL-10 engines. The advanced design performed by the orbit maneuvering pressure fed concept has been based on system. Gross attitude control during the pecformance potentially achievable these burns is provided by gimballing if an engine design could be developed i I the engines. All other orbital and for optim:j.m performance at both 100% reentry translational and attitude and 1% thrust levels. The current maneuvers are performed by the ReS. design high Pc engine performance is 14 MCDONNELL DOUGLAS ASTRONAUTICS If Report MOe E0056 Volume I 15 December 1969 estimated to be approximately 30 seconds RCS ENGINE ARRANGEMENT lower i~ Isp, which causes the pressure fed system to be 2000 pounds heavier ORBITE R - (U) 110 LB [NOIIES (4) 1&00 LB ENGINES than the RL-IO installation. Since the BOOSTER - (12) 110 LI ENGIIES IIOLB advanced design pressure fed and Lhe (4) 1&00 LUNGINES RL-IO concepts are essentially equal 110 LB ___~r----_--J (2) SHOIN FOR ORBITER in weight, the pressure fed concept } was selected for the baseline design to III REQUIRED FOR BOOSTER avoid the installation of additional engines. Figure 31 schematically shows the general arrangement of the pressure fed mode orbit maneuvering system. 1&00 LB Note that the propellant is drawn from IlO LB separate cryogenic storage tanks located in the aft section of the orbiter. Figure 32 ORBIT MANEUVER FEED SYSTEM A subsonic cruise propulsion sub system is incorporated on both the booster and the orbiter to provide the capability of (1) cruise back to the landing site (booster only) and/or landing assistance (booster and o~htter), LHl FLL DRAIN (2) go-around at the landing site, and / GOl VENT LOl FILL DRAIN __ ------~--;; RELIEF ------,- (3) cross-country ferrying. The cruise propulsion performance require ments for each of these operations are summarized in Figure 33. In addition, the study requirements were that only off-the-shelf engines using conventional r- - {x ENGINE-' \, JP fuel were to be considered in ISOLATION , detail. VALVE ---:------(~~~:/------~:::::::: ~,.. CRUISE PROPULSION REQUIREMENTS • CRUISE Figur. 31 • BOOSTER • RETURN TO LAUNCH SITE • RANGE CONTINGENCY - 20'~ • ENGINE-OUT - MAINTAIN 4000 FT MINIMUM ALTITUDE • ORBITER • NO CRUISE REQUIREMENT The number of Res engines and the • LANDING engine thrust levels may, for attitude • BOOSTER AND ORBITER control, be held to a minimum by • TOUCH-DOWN VELOCITY - LESS THAN 140 KHOn • ROLL-OUT - CONSISTENT WITH 8000 FT RUNWAY utilizing a combination of wing mounted • GO·ARDUND and fuselage mounted engines as shown • BOOSTER AND ORBITER in Figure 32. The translation engines • CLIMB RATE - GREATER THAN 2000 FT'MIN AT S.L. are also used for pitch and yaw attitude • PROPELLANT REQUIREMENT - 5 MIN AT TAKE-OFF POWER • ENGINE-OUT - GO-AROUND NOT REQUIRED control, with roll control provided by • FERRY additional wing mounted engines. • BOOSTER AND ORBITER Arrangements without wing mounting were • CLIMB RATE - GREATEIt THAN 400 FT 'MIN AT S.L. considered but would require additional • RANGE - GREATER THAN 400 MILES • AUXILIARY PROPULSION AND TANKAGE PERMISSIBLE engines or higher thrust levels to .NO PAYLOAD satisfy the yaw and roll requirements. • ENG INE -OUT - MAINTAIN 4000 FT MINIMUM AL TIT UDE Figur.33 15 NlCDONNIELL DOUGLAS ASTRONAUTICS Report MOe E0056 Volume I 15 December 1969 The baseline orbiter cruise pro efficient data management and display/ pulsion installation can be seen in control capability. It will relieve Figure 4. In this configuration four the crew of excessive workload by (4) JTBD-9 turbofan engines are automatically performing time critical mounted within the forward fuselage. functions and by providing priority sort The J~ fu~l is stowed in wing tankage. ing and data compression of that infor Doors are installed in each of the four mation needed by the crew. engines inlet ducts to protect the engines from boost and t:ntry heating. The general avionic functions are: The engine duct losses were estimated o Vehicle Self Test and Warning to be 5%. o Data Processing and Transfer o Crew Command and Integrated The engine exhaust ducts are Displays canted 20° to the vehicle axis. The o Target Tracking cosine losses were considered but o Autonomous Navigation and exhaust scrubbing losses on the side Flight Control of the vehicle were not evaluated. o Satellite Communications A detailed study of the effective o Supporting Energy Conditioning thrust loss aud the effects of noise and vibration induced on the sides of The key questions to be answered the orbiter should be accomplished in in order to define the Integrated future st'ldies. Avionics System are: the means of implementing Data Management; ~!-Board Unlike the orbiter, the booster Checkout; Display and Control; and has a long range cruise back require Reliability, as well as, Reuse. The ment. For this reason a significant features that were evaluated in pre portion of the system weight is fuel liminary tradeoffs in this study are and the operating duration of the indicated in Figure 34. engine will be hours instead of minutes. Thus the engine selection for the booster should hav~ the char acteristics of low specific fuel consumption rate and significant oper KEY QUESTIONS ating life. The baseline configuration DATA MANAGEMfjT • COMPIlTATIONS - DEc"nRALIZED uses six (6) JT3D-7 turbofan engines • INTERFACE TECHNIQUE IMUL TIPLEXEDI mounted in the forward fuselage as ON80ARD CHECKOUT • BUILT IN TEST shown in Figure 6. • LEVEL OF FAULT ISOLATION AND MAINTENANCE DISPLAY AND CONTROL • MUL TlfIIODE INTEGRATED IlISPLAYS • AUTOIIIA TIC SEQUENCING 7. Integrated Avionics - The RELIABILITY &REUSE • REOUNDANCY AND SELF TEST basic rationale for the use of Inte • MALFUNCTION DETECTION AND SwtTCHOVER grated Avion~.':s is derived from the • MAINTENANCE PHILOSOPHY measures required to achieve economy of operation. These measures are a self contained, crew controlled, pre Figule 34 launch che~kout capability, rapid turn around/reuse capability and a higher degree of mission suc~ess. Avionic The elements of the Integrated capabilities must include self check Avionics system are shown in Figure 35. out, block and functional reJundancy, Equipment and configuration selection and maintenance to a Line Replaceable was made on the basis of: (1) an es Unit (LRU). To ensure compatibility timate of the 1972 technology status with manned control, the Integrated and, (2) use of concepts which provide Avionics system will provide a highly small development risks. 16 ItIICDONNIEI..I.. DOUGLAS ASTRONAUTICS Report MOe E0056 Volume I 15 December 1969 BASELINE ORBITER INTEGRATED AVIONICS SYSTEM DISPLAY ARRANGEMENTS COWU~!c .. rlc~ - - 1 .-iH,-;-;-ll' ;;Hf ,1- T1It'''S((t\j' [illS ,INERTIAl ftASUREWNT IJ' • vOftTAC TltA"SO 1,,( II '1 • "AY COif"ut["IJ, • PII:OCEHO" 'J' .,UOU AlTI'(T[t\ : • 'INT(fltlf[O O,ftCAl" till 1 OUO • I"T[IIICOIII' "(lCU T\ 2 .l:·C1Un"S(·III) , • QIaI"-.T(ffJrelS I. • '''[''D[UCu~ "'OU! 1, • 'llwINCH! 'l ~ 1 • '1M' OISM ANTENN' J, • 'OOCII!"G S£'fSO"~ 1 OIU . ATTtluOI . -Ii' . \ ·(IIT, r~~~l-l ...... 'I.I .. T {OII'~l \ ,,,~(~ IOIT' OI\Plh" r':"lIilHt.:, 'Mlll fUCT~C'l • ,(fOE CDNTlI(h "IIIL ] .•-·J:~UllCUl-:ZOUT OJ.< • HAND CONT ~Oll E"~ ; • .• ue UN T \U"l'" l~ , • "til fl PI)'I'J>CSl .I!IITTUII[S11· CltT 1)1\PLI'\ ~ .INV("UIIS ,.1 • I'Ir AD UP OI~P~ lY i , • P.,NTfIt·l ~_ ~~"OU! I_~O!'.I_T~IMG J 1 ;: :~:~:A~~tn()tt iJI rllGMT cOfrlTJ', '''OCUSOlit ,}, 'O£NOns 5yHUIS NOT .p"1crm:.·' .INHIH".NUTI{)IIf UMO't~ uSltl 0Jt loosnlt • PO.{'t HIt~(I ~,,~ • I' It~ .1"¥OTt IillUL TIPl( 1(f'S 4 'rlt 'lMCT!O" • fUGHTJt[CQftOU!ll • .,H( G'''O ·!IoO' v" , I :"t.elT( h c... tllt .. , Figur. 35 Figure 36 Inertial sensors are used as the A common, multiplexed data bus prime source of navigation data through was selected to provide standardized all active mission phases. Choice of digital interfaces, and to reduce the inertial systems in both the booster complexity and weight of interconnecting and orbiter were dictated by the ascent systems. The intermix of computers guidance, entry to a pre-determined consists of a central data processor to landing site and automatic landing re perform mission oriented functions, and quirements. Star trackers and horizon peripheral dedicated cr.mputers for sensor sensors provide autonomous on-orbit functions, navigation, flight control, attitude and navigational updates. The and propulsion computations. This multi-mode rendezvous radar provides for arrangement was chosen on the basis of rendezvous with either cooperative or commonality of requirements while non-cooperat~"ve vehit~les. A dedicated maintaining equipment and software at navigation comptuer supplies the unique manageable complexity levels. Thus, requirements of individual system sensor oriented computational re sensors while permitting the central quirements, both hardware and software, sof~ware programming tasks to be main do not impact the central computer. tained at a manageable complexity level. On-board checkout minimizes ground This keeps sensor unique computational support and expedites maintenance and requirements from impacting the central reuse. Decentralized Built-In Test computational requirements. (BIT) was selected over a separate centralized test system to minimize The UHF communication link is interface complexity and provides sub utilized for EVA, inter-vehicle voice or system functional automomy.BIT pro data, and airport communication during vides self-test at all maintenance the approach and landing phase. The levels and permits identification of Comsat-link provides nearly continuous failures to the line replaceable units. communication capability between any Selective computar controlled areas ground station and the orbiter during for flight, caution and warning, or the orbital phase of flight. ground base checkout. The display concept utilizing Figure 37 shows size, power, and cathode ray tubes for "'Ilultimode data weight of the selected equipment. presentation permits crew d~cisions on Booster equipment is identical to that important tasks while relieving them of of the orbiter except that equipment the need to monitor a large number of utilized only for orbital operations displays and meters. A typical cockpit is deleted. Such equipment, as well display arrangement is shown in Figure as, the level of equipment redundancy, 36. is identified in Figure 35. 17 MCDONN.L~DOUGLA.A.7RONAUT'C. Report MOe E0056 Volume I 15 December 1969 ORBITER INTEGRATED AVIONICS that technology and research funding PHYSICAL CHARACTERISTICS is adequate to demonstrate feasibility of all technologies prior to go-ahead EQUIPMENT IEIGHT SIZE OPERATING POIER TYPE (LI) ICU FT) (IATTS) on Phase D. The development. manufac GUIDMCE & NAVIGATION 720 11.1 mo turing and flight tests efforts of this LANDING & NAVIGATlOI! AIDS 110 US schedule are considered to be the COIIIIUNICA TlONS us 41 .• •StS minimum allowable. The operati.,nal CENTULIiANACEII£NT COMPUTEII III 3.0 SOD program was assumed to ~Iave one launch DISPlAYS & CONTROL 411 US ISlS per month and require three orbital fLIGHT CONTIIOl 197 355 1115 ; vehicles and two boosters to meet this ! i CHECKOUT & MONITORING 125 2.1 260 ELECTRICAl 1160 170 10 III ICAP'CITY) schedule. Initial Operation Capability PIlI DISTil AND CONTROL .,R£ 700 10.0 (IOC) occurs in mid-J~ly 1976 and all SIGIlAL DISTil 1111£ IlOO 200 five production vehicles are utilized TOTAL ORIITEII AVIONICS 60St UU 5755 (P£AIII for flight testing. TOTAL BOOSTER AVIONICS 60.0 sou (PEAII) l. SUMMARY DEVELOPMENT SCHEDULE il figure 37 ",TIV,T, 1.1101 ""11171 "" 1j 8. Technology and Developr.lent Plan - This development test program provides a basis for establishing development costs. schedules, and identification of time critical develop ment effort where additional definition and study is required. A summary schedule illustrated in Figure 38 1s based on parallel develr~ment of the Orbiter and the Booster and assumes figllr.38 IJ Ii I iJ IJ 18 ItIICDONN.LL DOUGLAS ASTRONAUTICS Jl