NASA NACA-RM-E51A25-1951 Investigation of altitude ignition acceleration and steady-state operation with single combustor of J47 turbojet engine《带有J47涡轮喷气发动机的单一燃烧器高度点火 加速和稳态操作的研究》.pdf

《NASA NACA-RM-E51A25-1951 Investigation of altitude ignition acceleration and steady-state operation with single combustor of J47 turbojet engine《带有J47涡轮喷气发动机的单一燃烧器高度点火 加速和稳态操作的研究》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-RM-E51A25-1951 Investigation of altitude ignition acceleration and steady-state operation with single combustor of J47 turbojet engine《带有J47涡轮喷气发动机的单一燃烧器高度点火 加速和稳态操作的研究》.pdf（37页珍藏版）》请在麦多课文档分享上搜索。

1、LC N To RV L * F d I- I I ! I I I I t I I ! I RESEARCH MEMORANDUM INVESTIGATION OF ALTITUDE IGNITION, ACCELERATION AND STEADY-STATE OPERATION WITH SINGLE COMBUSTOR OF J47 TURBOJET ENGINE By Wiliiam P. Cook and Helmut F. Butee NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON March 5, 1951 Provi

2、ded by IHSNot for Resale-,-,-1 I N P 4 7 NACA RM E51A25 INVESTIGmION OF AUXTUDE IGNITION, AcCwIOm, AND STEADY-STATE OPERATION WITE SINGLE COWSTClR OF J47 TUR3OJEF EXGm By William P . Cook and H-t F. Butze SUMMARY An Investigation was .conducted qith 9 single combustor from a J47 turbojet engine usin

3、g weathered aviation gasoline and several spark-plug modifications to determine altituae ignition, acceleration, and steady- state opera5ing characteristics. Satisfactory ignition.was obtained with two modifications of the original opposite-polarity spark plug up to and including an altitude of 40,0

4、03 feet at conditiona simulating equilibrium windmilling of .the , engine at .a Zlight speed of 400 miles pr hour. ,At a sfmulated altitude of 30,000 feet, satfsfactory ignition was obtained over a range of sim- kted engine speeds. No significant effect of fuel temperature-on igni- tion limits was o

5、bserved mer a range of fuel temperatures frcsn 80 to -52O F. At an altitude of 30,OaO feet, the exces8 temperature rise avail- able for. accelemtion at low engine speeds was limited by the ability of the cambustor to produce temperature rise, whereaa at high engine speeds the zllaximum allawable tur

6、bine-inlet fxmperature became the re- atricting factor. c Altitude operattonal limits increased frm about 51,500 feet at 55 percent of rated engine speed to about 64,500 feet at 85 percent of rated speed. Cmbuetion efflciencies varied frm 59.0 to 92.6 percent over the range investigeted and deorease

7、d with a decrease in engine speed and with an increase in altitude; higher e$Picienciee would-have been obtained if lower altitudes had been investigated. Comparisons were made of the combustion efficiencies of weathered aviation gaeoline , Provided by IHSNot for Resale-,-,-2 NACA RM E5W5 and MIL-F-

8、5616 fie1 at altitudes of 30,000 and 40,000 feet. Canbustion efficiencies obtained with MIL-F-5616 fuel were B percent higher at rated engine speed and 14 percent lower at 55 percent of rated speed than those obtained with lieathered avFatiy, .and pressure drop, of single canbustors both of the annu

9、lar and of the can type have been investigated for different designs and for a number of different fuels (for example, references 1 to 4). Altitude ignition and acceler- ation a.m, of came, of great importanoe for multiengine planes having one or more engines temporarily inoperative or for aingle-en

10、gine fight- ers incurrfng blow-out at high altitudes. A study of the ignition charaoterietics of several fuels in a single can-type combustor is presented in reference 5 and a wind-tunnel inveetigatlon of altitude starting and acceleration oracteristics of the J47 engine is reported in reference 6.

11、In addition to such factors as inertia of the rotating parts and decreaaed air mass flaw at altitude, an important factor affecting acceleration of a turbojet plane is the temperature rise praduced by the ombustor in excess of that required to maintain the engine at steady-state operation for a give

12、n flight condition. This exoees temperature .rise available for acceleration is normally limited for two reasom: (1) Flame blow-uut may occur as the result of over-rich fuel-air ratios; or (2) allowable turbine-inlet tempraturee may be exceeded. The investigation reported herein was conducted t6 det

13、ermine the altitude ignition and acceleration characteristics of a single 547 cam- bustor. Additional data we=. obtained to evaluate the altitude oper- ational limits, ccanbustlon efficiency, and total-pressure losses of the combustor. Ignltim limits were determin+i at an altitude of 30,000 feet and

14、 at engine rotational speeds below and above equilibrium wind- milling speeds for simulated flight speeds of 400 and 354 miles per hour, respectively. Additional ignition-limit tests were made over a raw of altitudes for a simulated flight speed of 400 miles per hour and an en- gine speed equivalent

15、 to equilibrium windmifling speed. Acceleration . N 8 -4 -r Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS. XACA RM E51A25 3 cc characteristics were determined at a 30,OOO-foot sinerlated altitude over a wide range of engine rotational speeds (12.7- to

16、88.6-peroent rated engine speed) at a simulated flight speed of 400 miles per hour at and below equilibrium windmilling speed and 354 miles per harrr above equi- librium windmilling speed. All tests, including those for altftl.de operational limit and combustion efficiencyr, were made with weathered

17、 aviation gasoline that corresponded to MIL-F-5572, grade 115/145 fuel, from which 15 percent of the more volatile oonstituents had been re- moved to simulate altitude vaporization losses. Limited tests for comparisons were made with MIL-F-5616, a kerosene-type fuel that is the design fuel for the 5

18、47 combustor. . The installation of the 547 combustor photographicsally shm in figure 1 followed typical NACA procedure (reference 1) . A diagraslmatic sketch of the canplete experimental setup showing the location of con- trol equipent as well as the location of instmmntetion planes ie presented in

19、 ftgure 2. Instead of an electfic preheater, a gasoline- fired preheater (reference 4) was used. A detailed crosa-sect%onal sketch of the conibustor (including inlet and outlet diffusers having the same contour and dimensions as tne correspondfng engine parts) is sham in figure 3. me1 was supplied t

20、o the cambusor by meam of a duplex-type epray nozzle; the rate of fuel flaw was controlled by a manual valve looated downstream of a calibrated rotameter and a high- pressure pump and separated from the nozzle by apprgimately 10 feet of 3/8-inch outside-dimter tubing. Ignition was effected by =an8 o

21、f one of three different types of spark plug, a description of whioh follows. Plug A. - Two single electrodes of opposite polarity entereii frm diametrically opposed holes in the combustion chamber and formed a l/$-inch spark gap at the center line of the combustor, % inches frm the domed inlet end

22、(fig. 3). This plug, made at the Lewis laboratory according to the manufacturers recomendation, utilized most of the machined bodies of production pluss and had special poroelain insula- tors and center electrodes of 1/8-inh inside-diam=ter alloy tubing through which wae passed cooling air from the

23、combustor-inlet diffuser. Plug A was U6ed for most of the igni+ion tests and all other tests reported herein. 1 Plug B. - This plug was an experimental, opposite-polarity spark plug supplied by the manufacturer. The electrodes, instead of enter- ing from opposite sides of the ombustor, were about ll

24、Oo a and formed a 1/4-inch gap at the same position as plug A. The cooling alr Provided by IHSNot for Resale-,-,-4 - NACA RM E5lA25 for this plug entered through a 1/2 -inch hole in the skirt of the plug; the hole was located betweeh the combustor housing and the liner and faced upstream. The air en

25、tering the plug through this 1/2-inch hole then divided, a portion blanketing the exposed porcelain the re- - mainder passing down the hollow center of the electrode tubing (0.146-Fn. O.D.) and thence into the combustion chamber. Plug C. - This lug was the .%me as spark plug B except that the 1/2-in

26、cb air holes in the skirt of the plug were reduced to a 1/4-lncb diameter. A standard ignition coil supplied by the manufacturer was used in conjunction wlth all three spark plugs. Air flow and fuel flow to the combustor were metered by a standard A.S.M.E. thin-Ute orifice and by callbrated rotamete

27、rs, respectively. Temperatms and pressures of the inlet airwere measured by tWo eingle- junction irm-c;onetantan themooouples and by three, three-point total- pressure rakes ebnd one statio-presrJure tap, respectively, located at plane A4 (fig. 2) and 8rranged as shown in figure 4. Temperatures of t

28、he cmbuetor-exit gaaes were neasured by eeven banks of five-junction chromel-alumel thennocouple rakee looated at plane B-33 (fig. 2), corresponding approximately to the position of the turbine blades in the cmplete engine. Combustor-exit gas total pres- sures were measured at plane C -C (fig . 2 )

29、by Beven banks of five -point pressure rakes; a wall static.-pressure measurement was made at the aple plane. All total-preseure and temperature probes (fig. 4) were located at the centers of ewal areas, resulting in one pressure and me tem- perature reading for each 0.916 square inch of crose-secti

30、onal area. Fuel temperatws were measured by a single-junction lron-canertantan thermocouple in the fuel line immediately ahead of the combuetor. In order to investigate altitude ignition and aoceleratlan, it is necessary to determine the-engine oprating conditions that would be encountered at the su

31、dden opening of inlet.-alr gates to a parasitic- type engine at a given altitude and flight speed. Transient inlet- air conditions below equilibrium windmilling speed were o the fie1 flaw was then increased in 3 secosds to a value registered on the rotameter equal to three-fourths of the maximum fue

32、l-flow rate obtained vith slow-throttle advance. If flam blow- aut did not occur, the procedure was repeated to successively higher fuel-flaw rates. Canbustor-outlet temperatures attained after sta- bilization of combustion were used in the rapid-throttle tests be- cause of the difficulty In detemin

33、ing with thermocauples the inatan- tmeous average tempera.ture at the end of the 3-secod throttle etdvance. Altitude operational limits and combustion efficienoies at various shmlated-flight conditions were determined. Inlet -air temperature, pressure, and mass flow were set for- the partfcular flig

34、ht cordition temperature rise. Conditions at whiuh the required tenrpemture rfse “ investigated and the fuel flk adjusted to give the required combustor * Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS6 NACA RM E5M5 could not be attailzed were c-idered

35、to be in the inoperable range of the engine. All data were taken with the spark plug whereas at 40,000 feet,the allowable variation fram this fuel-air ratio was less than percent. 7 At a sFmulated altitude of 30,000 feet and a flight speed of 400 miles per hour, aatiafactory ignition w obtained with

36、 spark plug A at wtndmilling conditione ranging from 12.7 to 26.6 peroent of rated engine speed. At thls same altftude and a flight speed of 354 miles per hour, ignition was Battsfactory at steady-state conditions ran thus the reduced flow of cooling air to the electrodes served to improve the ignit

37、ion performEtnce of plug C over thaA of plug B. Acceleration The combustor tempemture rise obtained with a slow-throttle ad- vance is shown in figure 7 as a function of fuel-air ratio for various 50,000 feet. For the range of speeds investigated, from 15.2 to 63.3 percent.of rated engine speed, the

38、mimum obtainable temperature rise I windmilling and steady-state operating conditions at an altitude of - IS limited by b;lo+azt. In order to show more clearly the effect of increasing engfne speed, values of maximum obtainable temperature rise for both slow- and rapid-throttle advance at an altitud

39、e of 30,000 feet have been plot- ted in figure 8 as a finction of engine speed. Maximum-temperature- rise values for the slow-throttle advance were taken frcmnthe bluw-aut points shown in fi- 7; values for the rapid-throttle advance were obtained from the temperatures attained after s-hbilization of

40、 om- bustian. Rapid-throttle advance, as pviouely explained, canaisted . of a 3-second advance of the fuel tbottle to the maximum opening pos- sible without resultant blow-out. Also shown in figure 8 are the tempraturn rise required for eteadpstate operation of the engine at this altitude and flight

41、 speed (as calculated from the curves in fig . 5(b) ) and the maximum allarable temperature rise based 011 a limit - ing turbine -inlet temperature of 170O0 F. Provided by IHSNot for Resale-,-,-8 0 WA RM E5W5 With the rapid-throttle advance, the maximum obtaiuable tempera- ture rise increased from 8

42、80 F at about 32.5 percet of rated speed to 1500 F at 75 percent of rated speed. Fusrt;her temperature increases at engine speeds above 75 percent of rated speed were prohibited by excessive .cmbustor-outlet temperatures. With slaw-throttle advance, at conditions simulating steaQ-state engine operat

43、ion at a flight speed om milea per hour, the maximum obtainable temperature rise increase whereas at the high end of the speed range the ability of the themcmouples to with- stand the high ccxnlmstor-outlet temperatures required w thus, for the conditione investi at 55 percent of rated engine speed,

44、 however, operation with MIL-F-5616 fie1 resulted in a decrease of about 14 percent in ccanbustion efficiency. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS10 0 XMA RM E5lA25 Pressure Loss The total-pressure loss across the combustor is shown in figure

45、 12. The ratio of the total-pt?essure loss to a reference dynamic pressure AP/* is plotted .against the ratio of the IriLet-air density to the exhaust-ga6 derwrity and a straight-line correlation was obtained. Figure 12 indicates that the total-pressure loss ratio increased from 12.3 to about 16.1 a

46、s the dermity ratio increased from 1 to 2.6. From an investigation of the ignition, acceleration, and steady- state operational characteristics of a single oombuetor of a 547 engine using weathered ayiation gasoline with opposite-pdarity spark plugs at simulated flight conditions, the following resu

47、lts were obtained: 1. With combustor inlet-air conditione simulating equilibrium windmilling at a flight speed. af 400 pliles per hour, satisfactory ig- nitfon was obtained with spark plug A up .to .and .including 40,000 feet. No significant effect of -fuel temperature on ignition was observed over

48、the range of fuel temperatures investigate speea . speed At the same rotational speed at different flight Subtracting equation (15)- from equation .(16) yields (2 )VyrNz - (2 )Vz,Itz . - Assuming that the diffuser.efficiency is 100 percent and the velocity of the air leaving the diffuser is approxim

49、ately zero yields Y By substlution in equation (17), or Provided by IHSNot for Resale-,-,-3 NACA RM E5lA25 .- 17 Thus, if the combustor-inlet total pressure is known for an equilibrium windmilling speed at me flight speed, the inlet total pressure can be calculated for the same windmilling sped at another flight speed. Inlet temperature. - The total temperatures of the air at the combustor inlet for vmious engine Wrndmilling speeds