AERO PROPULSION EAS 4300
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This 14 page Class Notes was uploaded by Lauryn Sawayn DDS on Friday September 18, 2015. The Class Notes belongs to EAS 4300 at University of Florida taught by Staff in Fall. Since its upload, it has received 33 views. For similar materials see /class/206970/eas-4300-university-of-florida in Aerospace Engineering at University of Florida.
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Date Created: 09/18/15
Theory ofPropulsion PROPULSION EXERCISES B Equot p 20 o S 0 El 0 served into the 1960s This ghter which gured prominentl 39 James Michener s novel The Bridges of TokoRi was an improvement on the company s FH1 Phantom The aircraft used 2 Westinghouse J34WE30 turbojem delivering 3150 lbs of thrust eac 39 39 a m i speed of 587 mph and a range of almost 1300 miles The F2H could climb at about 7380 feet per minute and reach a service ceiling of48500 feet Ifthe airplane is operating at an altitude of 20000 feetwitha L 4 39 quot471 39 A7180in2 L 39r 39 p79psia nd the following a gm A r L 4 hv the pressure imbalance at the nozzle exit cexit weight ow rate if T71000F d ram drag at a ight velocity of 550 fps if wfwn3002 enet thrust f thmst power produced g exit velocity V7 h effective exhaust velocity Va ipropulsive ef ciency 71 i rm and r for HV18900 Btulb k the speci c fuel consumption s 2 The Redstone missile shown here launching the Freedom 7 spacecraft on May 5 1961 was the launch vehicle for America s rst man in suborbital space Alan Shepard This rocket can boost a 2850 lb 1300 kg payload to an altitude of 115 miles 185 km It is a one stage rocket vehicle using a single A6 engine with a burn time of approximately 155 seconds The propellant which is comprised of alcohol and LOX is consumed at the rate of about 300 lbss Experiments have shown that the exhaust velocity of the combustion products V78400 fps Assuming that the exit pressure of the nozzle is matched d termine a the thmst and thrust power under static conditions VDIO b the thrust and thrust power at a ight speed Vn5700 mph c the speci c impulse 19 d the overall efficiency hn at the given ight speed ifthe heating value of the propellants HV12800 btulb The heating value HV is typically quoted for the fuel alone and the quoted value of 12800 Btulb is the appropriate value for alcohol The optimal oxidizer to fuel ratio is r 147 for the alcoholLOX propellant combination 285 Theory 0 f Propulxion f an air aft whose total weight at the start of cruise rqu where WW IS the c bin w ight e powerplant and payl nd W is velocity Vr500mph and li to drag a39 the Breguet equation shows the range in cruise to be RLDWS 10 eWlW2 Assume that at the end of cruise all the fuel is used so that W2ng What is the effect ofsfc on the range the ratio erpWi is varied between 0 and 1 Vllhat has amore important effect on the cruise range sfc consider sfc between 05 and 15 lbs fuel per hour per pound of thrust or erpWi Substantiate your alyses conclusions with graphs andor sensitivity an 4 The 45 11 long BOMARC interceptor missile was powered by 39 en al LR59AG 13 liquid fuel booster rocket of 35875 lb 1596kN thrust and two Marquardt RJ43MA3 39 er ramjets of 11500 lb 512kN thrust each CM10A The maximum speed was 1975mph 3178kmhr and the ceiling was 550000 191mm with a r nominal inlet diameter of the en he is 24in 051m its length velocity of the engine vfzsoofps 762mS and assuming the ow conditions at e i e are e uivalent to those of night r e wer h the propulsive eflicien p i the ermal ef ciency up a the specific fuel consumption and k the overall ef ciency nu Theory of Propulsion 5 The Space Shuttle Transportation System SSTS incorporates 3 Space Shuttle Main Engines SSME using liquid hydrogen LHZ as the fuel and liquid oxygen LOX as the oxidizer in the ratio of 1 6 The V v heating value of LH2 is 51600 Btulb 120MJkg The thrust produced by each engine in the near vacuum conditions of the upper atmosphere is 512000 lbs 228MN at a specific impulse of 454 seconds The typical burn time for the engines is 85 minutes Determine the following a the effective exhaust velocity of the engine b the weight ow rate of fuel and of oxidizer for the engine c the total weight of the fuel and oxidizer required for the cluster of 3 engines d the volume for the fuel and oxidizer tanks e the length of the fuel and oxidizer tanks assuming they are cylinders equal to the diameter 276ft 84m of that of the 154 469m long Space Shuttle External Tank f the overall efficiency no at a ight speed of 6000fps 1 83kms mmr kl ll Inn 39x lam m I n l Halt ll nd mum Ulm 139er n H m ilhjll Dims rm lnlllm Immml quot nmss u 39 WI 17 l39qu39d hum nimm tum lbw qu liunl elullny Mnimm ENE Nun Tl tij Emmi MF WHHIW 39u lllmmul i39ugmwnmIwwmn39 Hum 287 Theory of Propulsion 6 Five Fl engines were used in the rst stage of the Saturn launch vehicle that lifted astronauts on a ight to the moon in the Apollo program The Fl still the most powerful rocket engine ever built uses RPl kerosene as the fuel and LOX as the oxidizer in the ratio of 127 The heating value of RPl is 18900 Btulb 44MJkg The tluu F1 ENGINE l st developed is l52x105 lbs 676MN at a speci c impulse of 265 seconds and u wi th a typical burn time is about 25 minutes Determine the following a the effective exhaust velocity of each engine b the weight ow rate of fuel and of oxidizer for the engine c the total weight of the fuel and oxidizer required for the cluster of 5 engines c the weight of the fuel and oxidizer required d the volume for the fuel and oxidizer tanks e the length of the fuel and oxidizer tanks assuming they are cylinders equal in diameter to that of the rst stage 33ft 10m f the overall ef ciency 1 0 at a ight speed of6000fps 183kms SPACECRAFI39 82 FT SATURN V LAUNCH VEHICLE 28 FL 1 Hlllllllll llllllllll UNIT FIRST ST SlC SHTUBN IJ LHUNCH IIEHICLE AGE mausr us um RSTAGE mam LBS NOTE WEIGHTS AND MEASURES GWEN ABOVE ARE FOR ME OMINAL VENICLE CDNFIGURAHON Fol APOLLO mi on HGLMES MAY VA SLIGHHY DUE to CHANG r MUNCH ro MEEI CHANGING commost w 288 Theory of Propulsion 7 The SCORPION Statoreaeteur Cruciforme eornrne ORgane Portant Integre39 a onliees Nasaux was a prototype of a long range ground to ai rnissile NIDZ6ange600 km o 220 mm liquid fueled rarnjet demonstrator designed by ONERA of Franee Successful ground tests of the fully integrated vehicle in ad Kerosene Flame eir 4 MA wind tunnel 1973 1974 air intakes tank holder 39 1 4 an 05 deterrnine the stagnation ternperature Tu leaving the eornhustor the percentage loss in stagnation ressure aeross R 116kJkgK and the average F133 8 The ASA F18 Hornet tem 12 a h 39 A n 39 l l 39 Dryden Flight Researeh Center Edwards California in a rnultiyear joint NASADODindustry prong the FA 1 p m h onverent nozzle like chat shown in due photograph is operated at standard 5 a level an Theory of Propulsion 7 t t D l 3amp2 9 Two 311771 alrcra have been used by NASA as testbeds for hlgirspe research The alrcratt an SRJIA and an SRJIB pllottralner alrcra hay Fllght Research Center Edwards Callfomla They were transfene to NASA a e U Alr Force ll 4 sonlc ow belng generated mg shape the ow can be d ed and hlghraltltude aeronautlcal e been ba e ASA s Dryden r super by the erhaust nozzles In a supersonlc nozzle whlch has a converglng dlverg acc elerate to supersonlc eed by passlng through the entlcal pom 1 at e Immum area f e yl 33 a rnlnlrnurn area Ag t z and an erpanslon out to an exrt area A7 at Whlch the pressure p7 matches the arnblent pressure pu5psla altltude27000 calculate a the yeloclty temperature pressure Mac A mum nl am statron c the goss num e thrust developed Weary amepuIsmn 45 degees super nun af 25 and sneexrsueseeum makmssnua afabvul7 percent Like me Wm uf 139s ancesmx me Feats me leading edge was equppea wnh summsue smsm m cant ed mcmpanud lam naps Thu reversal The heyemanated hunzanul m ablang nsse mm Shawn m mes gue pmveaes an immediate 2mgan resuxe uf me Hun senes af n me duct passes Ml 39 at 35mm almude Under ms sandman and sssummg me 1975 us mnde sunsspheee pmpemes apply deurmmz a me free mum staman pressue m b me the mem staguuan q q T m Lb mm W2 r assummg Lb gass ms m be me full a erburmng mm quated datemune me drag caef cxem cumgs when qxsthe dynme presslxe ands me wmg area Theory ofPropnlnon 11 223 at 40000 altttude lt wan weledb too GE J79GE17 llLIbOJElX ptoduotng 178801beolthroxt t each The mlEl med on the Phantom IS a quotD quot mlEl of anea 6 Sf mounted on etthen nde of the fuselage andcontatntng vanable angle nannpe and a boundary layer plttter plate to dwert lowepeedbonndary layer ow by the ramp euohthat 1t tnnptngee dnectly on the cowl hp The backplemme 1S adpntedeo down though a nonnal Sh ik wave an shown on the dtagrann and pm 1nto legton moquot and then on through the duct to the extt the NM tj the natto of the am at etattonZ g the rann dnag of the 1nlet F 292 Theory of Propulsion M M 20 pgo 435psia Tm 778R Inlet ramp angle 10 fuselage I Splitter plate Diverted boundary layer flow 12 Axial flow compressors are characterized by flow passages whose radial coordinates change little over the length of the machine while centrifugal compressors have blade passages that start near the axis of rotation and increase substantially in radius by the time the exit of the machine is reached Some idea of the difference is afforded by the accompanying photograph from NASA showing examples of two such machines side by side Consider standard sea level air entering a centrifugal flow compressor with the following conditions r2 6in c2 300fts and ocz 70 The air leaves the rotor with r3 18in c3 1200fts 0 25 The rotational speed of the compressor is 90001pm and it processes 3221bs of air Carry out the following a draw the combined entrance and exit velocity diagrams to scale b determine the entrance and exit blade angles 32 and 33 c determine the torque required to drive the compressor d determine the horsepower required e find the ideal pressure head f determine the contribution to the pressure rise provided by the external internal and centrifugal effects g determine the whirl velocity Ac h determine the ideal pressure ratio of the compressor Centrifugal compressor blade Centrifugal compressor blade passage Axial flow compressor blades Gap for stator passages 13 Repeat problem 12 changing only the entrance radius to r2r318in so that the machine now acts as an axial flow compressor 293 Theory of Propulsion 14 In 1965 GE won the contract to develop engines for the USAF39s Lockheed CSA 39Galaxy39 transport the first of the jumbo jets The GE TF39 engines introduced the high bypass turbofan Before the TF39 bypass ratios of turbofan engines were less than twotoone The TF39 showed that b ss ratios of 8 were possible far more than the bypass ratio of 2 that was common at the time This innovation resulted in specific fuel consumption as much as 25 less than previous other available engines In the late 6039s Lockheed and Douglas entered the commercial widebody airliner market with their trijet offerings GE introduced a new engine design the CF650 modeled after the TF39 for use on the Douglas DClO and Lockheed LlOll Considering a generic layout of the CF650 as shown below and assuming that the engine is operating in a groundbased test cell determine a the power developed by the low pressure and high pressure turbines b the thrust developed by the engine for the case of ideal expansion of the fan flow and the core flow c the specific fuel consumption Combustor Highp turbine Lowp turbine Theag amepulStan 15 The MQMVMA chmtar39 Brget rms had a may tapered ngeeshaped fuselage steetght mm H NDRYHRDP GRUMMAN Bum74E undEr the wmgs and a cunvmuunal tax cun gumuun wtth the tadplanes set m an mvened vee 1t ms guund ur a Shp Eng39 hamverisvjcs Nmnnn rpm Cumpressur type Centnrugat Turban type pr me E Axxal uw wtth e mtm em 2 Appueatmn Drunes Length n 57m Max Diameter n 27m Dry wetgqterengne 135kgAunewmte 1kg Turbmetmettempemture 95 c Theory ofPropulszon wcp um e n t an I mun 39 us Yi lucgusnm a LINE EMquot igsnanl hem r Manon m r r Y n5 Fur sea1eve1 stahe enhmhnhs and assummg pupfo 9 45 40 40 and HV44MJkg In th mm t m hm an PM at thrust at sea level The eharaetenstaes er the enmhustmh prnduets are enmhustmh eharh er temperature 530011 average wem heat ratm F1 2 mnlecular werght W 7 and chamber pressure pl 34 atmnspheres d t t t 77 95 b the prnpeuant nw r t 5km a fur thrs amtude furmatched nnzzle exrt nperatznh 17 en Delta launchers 1t 15 an 1t 15 an RFrlLOX ehgme that develnps an 12t34s at sea level and an We 7 99 see m space and hr Luam er r Exansmn mun F10 The ehgrhe wer hs apprnxrrhately an ratm M e heats are Ennsldered Theory amepuIsmn V mamy cae xcxems are in us and of n m mmng Under these sandman and assuming sea u m charactmsmc mamy 0 13 9mm atsealevel The chambexpxessurep l amasphzxesand39he cambumanpzannc39s have the m V n arp T M mav slang the nmzle whase caaxdlmtes are gven m the um belww m terms af anal mum we be Jaunted Uang he rm dimensanal results 515a draw the nmzle Ixm liz nlrzn 45 Ha Ira In In mu m In ml lz s 25 m m In I In m m In a m mm mm a nmapmxe mm mm mu T uummmwaq expanem nn 35 pmpenm denmy p mamm3 temperature cae xcxem af pressure n 1 we charactmsmc exhaust mamy wannas cambmhan tampexmxe TK6EI92F effecuve m 3 my 147 393 E E WmaWWW WWW a m m g m M m M mm m M snmmm zwmgaan c m umpmm af T IS 39 at a manual m x 155 The characunshcs af the pmpgumg a patasmxn sea level mm 491 and a bun hm uf expussedm amasphexes Eummgnte cms Hm l zexpm nmsarxsmpm Chuacunmc velacxtycms cquot muman EIEIEIZITJSpquotm5 Specx chutnua H 27 u wexght wan Pmpellmtdemtyp 177nkgm3 Chamber press Wu 5 atm Theory of Propulsion Determine a the propellant grain dimensions A and L b the specific impulse c the chamber temperatured the in uence of a 1 variation in the buIning rate of the propellant on the thrust and the chamber pressure e the motor performance at TFSSC and 25C 298
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