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Research Memorandum RESEARCH MEMORANDUM ALTITUDE OPERATIONAL CKARACTERISTICS OF A PKOTOTYPE MODEL OF THE J47D (RX1-1 AND RX1-3) TURBOJET ENGINES WITH INTEGRATED ELECTRONIC CONTROL By E. William Conrad, Harry E. Bloomer, and Adam E. Sobolewski NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON January 8, 1952 .. .. .. I ,.. ... .. .. .. CINCLASSIFIED 1N NACA RM E5U?,O8 L U It w0 ;j MODEL CIE' ICBE J47D (RX1-1 AND RXl-3) TURBOJEIT By E. William Conrad, Ea- E. Bloomer, and Adam E. Sobolew8ki SUMMARY The altitude operatianal characteristicsof a prototype modelof the J47D (RXl-1 and RXl-3) turbojet engines, which includesan after- burner, a variable-area exhaust nozzle, anand integrated electronic control, were investigatedin the NACA Lewis altitude wind tunnel at altitudes from10,000 to 55,000 feet ata flight Mach numberof 0.19 and at fli@t Mach numbers from 0.19 to 0.89 at as altitude of 25,000 feet. Data obtained with oscillograph recorders and conven- tioa instrumentation are presented to showfollowing the charac- teristics: (a) Compressor stan cb) Combustor blow-out (c) Acceleration (d) Deceleration (e> Altitude starting [f) Afterburner ignition (g) Afterburner transients (h) Afterburner blow-out For both of the engines investigated(FZL-1 and RIEL-31, it was found that the compressor-s€alldata plotted as single curveson coordinates of compressor pressure ratioand corrected engine speed. TKO distbct types of stall appear to exist with the transition occur- ring at corrected engine speeds between5250. and 6250 rpn. The com- pressor unstall characteristics areshown on the same coordinates. 2 NACA F44 E51Eo8 rpm, For corrected engine speeds above5800 the pressure ratlo neces- +L sary to unstall the capressor lower was than %he pressure ratio for either stall or steady-state operation.As the altitude wss increased, the compressor unstalled at slightly higher pressure ratios; however, flight Mach nuniber had no apparent effect within the range investigated. P Combustorblow-out data for .dlflight conditions investigated with 8 the RXL-1 engine were plotted aas single curveon the same coordinates and coincided almost exactly with the compressorstall curve. Acceleration time from the toidle the rated thrust condition without afterburning Fncressed from14 to 22 seconds as the altitude was increased from 15,000 to 45,000 feet ata flight Mach number of 0.19. At an altitude of 25,000 feet, &~1increase in flight Mach number from0.19 to 0.75 reduced the accelerationtime from idle to rated thrust without afterburningfrom 14.4 to 6.0 skconds. When a minlmum fuel limit of 450 jpunds per hour ms used, lean combustor blow-out could not be obtainedduring deceleration for the range of flight conditions covered by the investigation. ’ Ignition in all combustors was obtained at.windmilling speeds from ,1300 to 1500 rpu at an altitudeof 50,000 feet usingMIL-F-5624 (AN-F-58) fuel ata temperahre of about 70° F and inlet-air tempera- tures ofOo to -So F. At 35,000 feet, ignition occurredFn all com- c bustors up to,3500 rpm,. the highest windmilling speed obtainable. With the useof MIL-F-5624 fie1 (treatedto give a l-pound Reidvapm pressure) ata temperature of apprdximately90° F and inlet-air tem- peratures from-200 to 300 F, ignition was possible inall combustors at 49,000 feet .upto a windmilling speed .of about JE.0rpn. At 25,000 feet, however,.starts were possible upto 2200 rpn. At an altitude of 40,000 feet, theoptimum fuel flow for starting appeared to be about650 pounds. per hour for ME-F-5624 f’uel with a 7-pound vapor pressure. merburner starts by autoignitionwere obtained at altitudes up to 53,000 feet at a flight Mach numberof 0.19 using MIL-F-5624 fuel. The tail-pipe fuel-air.ratios required for autoignition increased with.. altitude d.at35,000 feet decreasedas the burner-inlet temperature was raised. The tail-pipe _fuel-air ratio at leanwhich blow-out of the afterburner occurredwas increased as altitude was raised. The width of. the blow-out- band- also .increase& with altitude. RACA RM E5lXO8 3 The improvement in performance characteristicsof turbojet engines effectedby the applicationof a$terburning and a continuously variable-area exhaust nozzle have been well established during the past few years. Manual control of a turbojet engine equipped withan afterburner and variable-area exhaust nozzlemad, however, placea heavy burdenon the pilot or flight engineer.To relieve the pilotof the extreme complexityof engine operation and 'the need for constant surveillance of mahy engine operation limits,a completely automatic control system is required. Accordingly, a prototype modelof the J47D (RXI-1 and RXL-3) tur- bojet engineswas provided withan afterburner, a variable-area nozzle asd an integrated electronic control. The enginewas installed in the NACA Lwd-6 altitude wind tunnel to obtain the performance character- istics and insight into the engine's operational problems. The steady- state performance of the engine without- afterburning isin presented references 1 and 2. me operational characteristics are presented herein. The integrated electronic control usedthe with prototype model of .. the J47D (RXl-1 and Rx1-3) turbojet engineswas designed to (I) provide - s-e-lever thrust control over the full range of operation from starting tofull afterburning condition, (2) schedule all services required, and (3) give the maximum acceleration and deceleration rates I possible without exceeding engine speed and temperature limits or causing cdustor blow-out or compressorstall. To attain the latter objective, the ccmgressor stall and combustor blow-out regions. were investigatedwith the control inoperative. The limits maintained by I the control were then adjustedas necessary and the transient per- formance of the controlled engine was evaluated. Oscillograph traces axe presented hereinto show the typical behavior of these variables during compressor stall, comp1et.e and partial conibustor blow-out during acceleration, controlled accelera- tion and deceleration, and afterburner ignition. Compressorstall and combustor blow-out limits have been correlatedto show the effects of altitude, flight.Mach numberand corrected,engine speed. The effects of altitudeand flight Mach number on engine acceleration snd deceleration arealso given, as wellas steady-state operational char-. acteristics suchas engine and afterburner ignition apd aFterburner 4 NACA HM E5lEO8 lean blow-out. Data were obtained at altitudes from 10,000 to 55,000 feet at a flight Mach.number of 0.19 and at flight Machnumbers from .. 0.19 to 0.89 at an altitude of 25,000 feet: . DESCRIPTION OF ENGIXE The J47D RXl-1 and RXl-3 engines used in this invesfiigation were aerr>"xLly the same as the J47D engine. - The manufacturer guaranteed static sw-level performaace of.tk J47D engine is 5670 pounds thrust at 7950 rpm'ana an exhaust-gas telqperature of 1275O F. The main chpnents include a 12-etage axial-flow compressor with a pressure ratio of about 5.1 at rated conditions, eight cylin- .. drical direct-flow combustors, a single-stage turbine, a diffuser, an afterburner maustion cbber, a variable-&ea exhaust nozzle, and an integrited electronic control. The over-all length of the engine including the afterburner is about 217 inches', the maximum diameter is approximately 37 inches, and the total weight is about 3000 pounds. A view of the turbojet engine installed in the test section of the altitude .wind tunnel is given in fLgure 1 and a schematic drawing of the engine ie presented in .figure 2. Two combustor configurations were used with the RXL-1 engine. The original configuratfon included ,conventional-type spark plugs,. a 20,000-V0lt, coil and vibrator unit, and cross-flre tubes I&inches in diameter. The second configuration,referred to as the Iff modified combustor," had cross-fire tubes that were 2 inches in diameter. The modified combustor liners also had semlcylindrical shroud8 projecting 1/4 inch radially inward from the damstream half of each of the first six rows of holes. These ch&nges,- as well a8 other ininor hifferences between the original and modified combustor liners are shown in figure 3. The ignition systems used with the modified configuration were (1) qpposite polarity spark plugs mounted as shown in figwe 4 'and (2) an ignition system which produced a potential difference approximately twice that of the standaxd system. Only the modified configuration was used with the Rx1-3 eane. For both combustor configurations, two sets of spark plugs, located fa diametrically opposite combustors, were. used, . .. The afterburner shown schematically in figure 2 wa6 camprised of a diffuser 43 inches in length, a ccmibustionchamber 50 inches in length ' which tapered from a J%-inch1 diameter at the flame holder to a 29-inch diameter at the e&ust-nozzle section, and a variable-area exhaust nozzle whfch was 16 inches in Length in the open position. Flame-holding surfaces were provided- with 2-rbig, V-gutter flame PTllCA RM E51EO8 5 holders used in conjunction with a d-inch diameter pilot cone. The 2 location of the flame holders was varied from 2 to 7 inches downstream of the pilot wne. Fuel was injected by means of 24 radial-spray bars (two sets.of 12) located and 13 inches upstream of thepilot cone. The function of the integrated electronic control system is to cause the engine to operate at an optimum point determined from performance, operational, and safety considerations at asy given operating condition. - A compreheneive study of the control is given in reference 5. A block diagram from this reference is given in figure 5 to show therelation of the control components. Detailed - functions, some of which are discussed more fully in appendix A, my be listed as follows: 1. Maintain a-set enginespeed irrespective of changes in alti- tude or flight Mach nmber and maintain rated engine speed within close limits during steady-state operation 2.
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