nL-ft ^f^yC! 1. RM E58C12 I— I o 00 H o <; 2: NACA RESEARCH MEMORANDUM PERFORMANCE OF BASIC XJ79-GE-1 TURBOJET ENGINE AND ITS COMPONENTS By Carl E. Campbell Lewis Flight Propulsion Laboratory Cleveland, Ohio GAINESVILLE FL ^o^^^ -. ^'i-L 32611-7011 USA NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON May 8, 1958 (Second printing, for non -military distribution, May ^1, I961) jil / 'T'^ 77 yfT^f^y NACA RM E58C12 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS RESEARCH MEMORANDUM PERFORMANCE OF BASIC XJ79-GE-1 TURBOJET ENGINE AND ITS COMPONENTS By Carl E. Csmpbell SUMMARY An investigation to determine the performance of the XJ79-GE-I tur- bojet engine and its components, while operating as integral parts of the engine, was conducted in an altitude test chamber at the NACA Lewis lab- oratory. Data were obtained over a range of Reynolds number indices from 0.60 to 0.08 and for various settings of the variable compressor stators and variable-area exhaust nozzle from fully open to fully closed positions. Compressor performance and turbine performance are presented in the form of performance maps at selected values of Reynolds number index; the effects of Reynolds nimiber on performance are summarized. The ef- fects of variable stator angle and high inlet-air temperatures on com- pressor performance are also shown. Combustor performance is given in generalized form as a function of the usual combustor parameters. Ex- haust system data are presented to permit the calculation of over-all engine performance from pimiping characteristics. Maps of engine pumping characteristics are presented at selected values of Reynolds number index, and the general effect of Reynolds number on the pumping charac- teristics is summarized. Over-all engin^ performance (net thrust and specific fuel consumption) is presented for a flight Mach nvmiber of 0.9 at rated engine conditions over a range of altitudes to illustrate per- formance losses resulting from decreased Reynolds number index. All component and engine performance data are presented in tabular as well as graphical form. INTRODUCTION An Investigation to determine the performance of the XJ79-GE-1 tur- bojet engine and its components while operating as integral parts of the engine was conducted over a range of Reynolds number indices in an alti- tude test chamber at the NACA Lewis laboratory. This engine incorporates variable inlet-guide vanes and variable stator vanes in the first six compressor stages as a means of avoiding part-speed surge. The engine also has an afterburner and iris-type variable primary and secondary NACA RM E58C12 exhaiost nozzles; however, during the investigation reported herein, the afterburner was inoperative and the secondary nozzle was removed. The variable-stator system and the variable exhaust nozzle are normally scheduled automatically by a combination electronic and hydraulic con- trol, but they were manually positioned during this investigation. The performance data were obtained over a range of Reynolds number indices from 0.60 to 0.08 with the variable stators in the open position and over a range of stator positions from 0° to 35° at a Reynolds number index of 0.20. At each Reynolds number index and stator position, data were obtained over a range of engine speeds at five exhaust -nozzle areas. Data were obtained at engine inlet temperatures up to 700° R, but the bulk of the data was obtained at an inlet temperature of approximately 416° R. Component performance data are presented over a range of Reynolds number indices and variable stator positions. Generalized engine data are presented in a form that permits computation of engine performance at operating conditions other than those specifically investigated, and the method of such computation is illustrated. All component and engine performance data obtained during this investigation are presented in tabular as well as graphical form. APPARATUS Engine The XJ79-GE-1 turbojet engine has a length of 207 inches and a maximimi diameter of 32.6 inches at the turbine section. The frontal area based on the compressor tip diameter is 4.89 square feet. The dry weight of the engine and its accessories is about 3150 pounds. The maniif actur er ' s static sea-level military performance rating (nonafter- burning) is 9600 pounds of thrust with a specific fuel consumption of 0.87 pound per hour per pound of thrust at an engine speed of 7460 rpm and a turbine -out let temperatiire of 1070° F. The XJ79-GE-1 turbojet engine has several minor airflow bleeds that are used for cabin pressurizatioh, anti-icing, tiirbine cooling, and bearing-seal pressurization. These bleed flows are extracted from the main engine airflow at the seventh and ninth compressor stages and at the seventeenth compressor -stage seal. The amount of bleed flows dumped overboard during this investigation did not exceed approximately 1.5 percent of the inlet airflow; the remainder of the bleed flows reentered the main stream before reaching the afterburner diffuser section. NACA EM E58C12 Engine Components Compressor . - The seventeen-stage axial-flow compressor has variahle inlet guide vanes and variable stator blades in the first six stages that are moved simultaneously from the open position to their respective closed positions. The angle of travel from open to closed for the vari- our stages is as follows: inlet guide vanes, 47°; first stator stage, 44°; second through fifth stages, 35°; and sixth stage, 41°. All ref- erences to stator position throughout the report will be in terms of the second through fifth stages as a matter of convenience; that is, the closed position will be referred to as 35°. The compressor has a con- stant tip diameter of 29.95 inches through the first fourteen stages and tapers down to a tip diameter of 29.3 inches at the seventeenth stage. The hub -tip radius ratios of the first, fourteenth, and seventeenth stages are 0.36, 0.86, and 0.88, respectively. The compressor was de- signed to deliver an airflow of 162 pounds per second and a total- pressure ratio of 12.2 at static sea-level military conditions. Combustor . - The combust or is a cannular type with ten circular through-flow inner liners. Fuel is supplied to each liner through a single-inlet duplex fuel nozzle. Ignition is provided by a spark plug in one of the inner liners and spreads to the other liners through interconnecting crossfire tubes. The combustor-inlet reference velocity, based on the full b\arner section area of 4.33 square feet, is approxi- mately 89 feet per second at design sea-level conditions. Turb ine . - The three-stage impulse -type turbine has a constant pitch-line diameter on all three stages and tip diameters of 28.4, 29.65, and 31.05 inches for the first, second, and third stages, respectively. The hub-tip radius ratios of the first, second, and third stages are 0.81, 0.70, and 0.59, respectively. The increase in annular area through the turbine occurs entirely in the turbine nozzles. The t-urbine was de- signed to operate at a turbine-inlet temperature of 1700° F at 7460 rpm. Damping rods are installed between adjacent third-stage rotor blades to reduce blade vibration. These rods are 3/16 inch in diameter and are situated at approximately 75 percent of the blade height. Control and exhaust system . - The engine control schedules the variable -stator assembly to vary continuously as a function of corrected engine speed from the closed position (35°) at 64 percent of rated cor- rected speed to the open position (0°) at 90 percent of rated corrected speed (fig. l). During the investigation, the original schedule was altered as shown by the dashed line of figure 1 to achieve a higher thrust during operation at a high engine-inlet temperature corresponding to a flight Mach number of 2.0. The primary exhaust nozzle, which is a convergent, variable-area iris-type nozzle, is scheduled to vary in grad\aal steps from an open position at idle conditions to a closed NACA RM E58C12 position at military position of the power lever. During this investi- gation, however, the variable-stator assembly and the exhaust nozzle were manually controlled. The afterburner, which was inoperative during this investigation, has a maximiim internal diameter of about 34 inches and includes a diffuser, fuel-injection bars, a pilot burner, a three- ring gutter -type flameholder, and a corrugated and louvered cooling liner. The variable secondary exhaust nozzle was removed during this investigation. Installation A view of the XJ79-GE-1 t^urbojet engine installed in the altitude test chamber is shown in figure 2. The engine was rigidly mounted on a flexure-plate supported test platform that was connected by a linkage to a calibrated null-type thrust cell. Dry refrigerated or heated air entered the engine inlet through a bellmouth Ventiiri duct, which was mounted to the engine inlet and test platform. The inlet section is separated from the exhaust section by the front bulkhead, which incor- porates a labyrinth seal around the inlet Venturi duct to prevent the flow of combustion air directly into the exhaust section and to permit the measvirement of thrust forces. The inlet- and exhaust -pressure con- trols are designed to maintain automatically a constant ram-pressure ratio and exhaust pressure. Instrumentation Instrumentation for measuring pressures and temperat-ures was in- stalled at various stations through the engine as shown in figure 5. The table presented on the figure indicates the number and type of meas- urements at each station. Total-pressure and temperature probes at each station were located at the approximate centers of equal annular-area increments so that measurements could be averaged arithmetically. In- strumentation was also provided to measure the portion of bleed flows dumped overboard through the compressor-discharge standpipes. Pressiires were measured by null-type diaphragm capsules and recorded by a digital, automatic multiple -pressure recorder. Temperatures were measured and recorded by iron-constantan and Chrome 1-Alumel thermocouples in conjunction with self -balancing potentiometers. Fuel flow was meas- ■ured by a calibrated turbine-type flowmeter. The variable stator posi- tion and primary exhaust -nozzle area were determined from cold calibra- tions of output voltages from linear potentiometers in their actuating mechanisms. NACA RM E58C12 PROCEDURE Most of the data were obtained at the miaimuin inlet -air temperat'ure consistently available (approximately 416° R) to extend the range of cor- rected engine speeds. Engine-inlet pressures were selected in conjunc- tion with this inlet temperatin'e to give a Reynolds number index range from 0.60 to 0.08. Some data were obtained at higher inlet-air tempera- tures up to 700 R at a constant Reynolds number index of 0.4 in order to investigate the effect of temperature itself on the reproducibility of the data. With the variable stators in the open position, data were obtained at each Reynolds number index at five fixed settings of the variable- area exhaust nozzle over an engine speed range from military (7460 rpm) down to the surge region. At a Reynolds number index of 0.2, similar data were obtained for other settings of the variable stators down to the fully closed position. Fuel conforming to the specification MIL-F- 5624A grade JP-4 with a lower heating value of 18,700 Btu per pound and a hydrogen-carbon ratio of 0.171 was used throughout the investigation. Definitions of symbols, methods of calculation, and a sample calculation of engine performance from generalized performance data are presented in appendixes A, B, and C, respectively. RESULTS AND DISCUSSION Component Performance Compressor performance and turbine performance are presented in the form of performance maps at selected values of Reynolds number index, and the effects of Reynolds number on performance are summarized. The effects of variable stator angle and hot inlet -air temperat\ires on com- pressor performance are also shown. Combustor performance is presented in a generalized form as a function of the usual combustor parameters. Exhaust -system data are also shown to permit the calculation of over-all engine performance. Compressor performance . - The compressor performance map at a Reynolds number index of 0.60 with the variable stators in the open posi- tion is shown in figure 4(a). At design compressor pressiire ratio (l2.2) and rated corrected engine speed (7460 rpm), the corrected airflow was approximately 159 pounds per second, and the compressor efficiency was 0.784. Compressor efficiency reached a maximum of 0.80 to 0.81 at a corrected engine speed of approximately 6900 rpm. At a given corrected engine speed, variation in compressor pressure ratio (as limited by operation of the compressor and turbine as engine components) caused variations in compressor efficiency only of the order of 0.01. The cor- rected airflow was unaffected by coirrpressor-pressiire-ratio variations at NACA RM E58C12 high corrected engine speeds; but at a corrected speed of 6600 rpm, the corrected airflow was lowered about 3 percent by increasing pressure ratio over the range permitted by the variable-area exhaust nozzle. The open-stator compressor performance map at a Rejmolds number index of 0.20 is shown in figure 4(b). At the design pressure ratio and rated corrected engine speed, corrected airflow was about 2 percent lower and compressor efficiency was approximately 0.02 lower than at a Reynolds number index of 0.60. Peak compressor efficiency was approximately 0.03 lower and occurred at a slightly higher corrected speed. Variations in pressure ratio at a given corrected speed had a greater effect on cor- rected airflow at this lower Reynolds number index (0.20). At a cor- rected speed of 6600 rpm, increasing the compressor pressure ratio over the permissible range lowered the corrected airflow about 6.5 percent at a Reynolds number index of 0.2 compared with a 3 percent reduction at a Reynolds number index of 0„6. The comtpressor performance at a Reynolds number index of 0.20 with the variable stators fully closed is shown in figure 4(c). The variable stators are scheduled to be closed only for the low-speed portion of the map. The compressor performance was mapped over the full range of cor- rected engine speed and exhaust-nozzle area at this stator position (fully closed) to permit the determination of compressor performance for other possible schedules of engine speed with stator position and other exhaust-nozzle area schedules. With the stators closed, the compressor operates in a region of considerably reduced airflow and pressure ratio and the peak compressor efficiency is less than 0.60. Corrected airflow was sensitive to compressor pressure ratio over the entire corrected engine speed range. The slope of the corrected speed line at 5200 rpm illustrates the advisability of scheduling a large exhaust -nozzle area at this operating condition. Compressor efficiency and corrected airflow for open stator opera- tion are shown in figure 5 as functions of Reynolds number index for constant values of corrected engine speed and compressor pressure ratio. Reynolds number variations had no appreciable effect on compressor effi- ciency or corrected airflow for values of Reynolds number index greater than approximately 0.4. At a corrected engine speed of 8000 rpm and a pressure ratio of 12.50, reducing the Reynolds number index from 0.4 to 0.08 lowered the compressor efficiency by 0.05 and lowered the corrected airflow approximately 7 percent. The decrease in efficiency and cor- rected airflow with decreasing Reynolds number index was greater at the lower corrected engine speeds. Compressor performance losses due to Reynolds n\amber effects were also greater at higher compressor pressure ratios, especially in the low corrected-engine-speed region. NACA EM E58C12 Corrected airflow data obtained over a range of iolet-air tempera- tiires corresponding to flight Mach numbers up to 2.0 did not appear to generalize within the expected acciiracy of the airflow measurements (fig. 6(a)). The apparent trend of increasing corrected airflow with increased inlet temperature was still doubtful because the compressor pressure ratio was not quite constant in spite of the constant exhaust- nozzle area^ and the hot -temperature data were in the low corrected- engine-speed region where pressiire ratio was shown to have an effect (see fig. 5). However, after the effects of compressor pressure ratio were eliminated (fig. 6(b)), the trend of increasing corrected airflow still existed and amounted to approximately a 3 percent increase in corrected airflow for an increase in inlet temperature from 416° to 700° R. Although the reasons for this effect are not understood, it is believed that the accuracy of the data involved establishes the existence of this trend. The effect of stator position on compressor performance is shown in figure 7 for a range of exhaust -nozzle areas at a Reynolds number index of 0.20. Corrected airflow with the variable stators closed was approxi- mately 38 percent of the open stator value at a corrected engine speed of 8000 rpm and about 44 percent of the open stator value at a corrected engine speed of 6700 rpm (fig. 7(a)). The corrected engine speeds at each stator position that correspond to the control schedule values (fig. 1) are shown on the figure by vertical dashes. The corrected airflow for off-schedule operation or for altered schedules may be approximated by linear interpolation. Compressor pressure ratio (fig. 7(b)) with the variable stators closed was about 36 percent of the open stator value at a corrected engine speed of 8000 rpm and about 42 percent of the open stator value at 6700 rpm. At the corrected speed of 6700 rpm, the closed nozzle pres- sure ratio of 9.1 could be reduced to 8.3 by opening the exhaust nozzle or to 3.85 by closing the stator vanes. The low-speed stall- line inter- cept with open stators and an exhaust-nozzle area of 2.81 square feet and the approximate scheduled operating line of the variable stators are indicated on figure 7(b) to illustrate the necessity of an antistall device on this high-conipression-ratio compressor. The effect of stator position on compressor efficiency is shown in figure 7(c). Peak compressor efficiency was lowered from 0.77 to 0.75 by closing the variable stators from 0° to 18°, but dropped off rapidly to less than 0.60 when the stators were fully closed to 35°. The cor- rected engine speed at which the engine is scheduled to operate for a given stator position was in the region of peak compressor efficiency for the two closed-stator positions investigated. NACA RM E58C12 Combustor performance . - The variation of combustion efficiency with the combustion parameter w g'^'g is shown in figure 8. This parameter is approximately proportional to the combustor-inlet parameter PT/V and is more convenient in conjunction with over-all engine performance calculations. Combustion efficiency varied from 0.97 at the highest values of w^^ g^g to 0.90 at a combustion parameter of approximately 15,000, which essentially covered the range of engine and flight condi- tions investigated with the variable stators in the open position. With the stators closed at a Reynolds number index of 0.20, the combustion parameter could be reduced further, and the combustion efficiency dropped off rapidly with decreasing engine speed and increasing exhaust -nozzle area to about 0.33 at a combustion parameter w^^ g'^g °^ 4500. However, this condition would not normally be encountered in actual flight and is presented as isolated combustor performance beyond the usual range of a combustor operating as an integral part of the engine. The combustor- inlet conditions of pressure, temperature, and reference velocity at this point were approximately 750 pounds per square foot absolute, 610° R, and 60 to 70 feet per second, respectively. The corresponding inlet con- ditions for the maximum combustion parameter and combustion efficiency were 13,626 pounds per square foot absolute, 1078° R, and 82 feet per second, respectively. The combustor total-pressTire loss ratio as a function of combustor temperatiire ratio is shown in figure 9. As the temperature ratio in- creased from 1,45 to 2.05, the combustor total-pressure loss ratio de- creased from 0.07 to 0.05. This reduction in total-pressure loss ratio results from the more rapid decrease in friction pressure loss that ac- companies the decrease in combustor-inlet Mach number in comparison with the increasing momentum pressure loss as combustor temperature ratio is increased. Turbine performance . - The over-all performance of the turbine is presented in terms of corrected turbine enthalpy drop and turbine gas- flow parameter for lines of constant corrected turbine speed, pressure ratio, and efficiency. The turbine performance map determined from open stator data at a compressor-inlet Reynolds number index of 0.60 is shown in figure 10(a). The range of engine speeds and exhaust -nozzle areas investigated caused the turbine -inlet Reynolds number index to vary from 1.04 to 1.41, but, as will be shown later, Reynolds number has little or no effect on turbine performance in this range. Over the narrow range of corrected turbine speed and pressure ratio as limited by operating in an engine at a constant stator position, the turbine effi- ciency varied from 0.86 to 0.88. The corrected turbine gas flow, which can be obtained by factoring out the corrected turbine speed and the factor 60 from the turbine gas-flow parameter, was about 28.3 pounds per second. NACA RM E58C12 At a compressor-inlet Reynolds number index of 0.20 and open stator position, the Reynolds number index at the turbine inlet varied from 0.29 to 0.43 (fig. 10(b)). The turbine performance map was quite similar to that at the higher Reynolds number index; the corrected turbine gas flow was still about 28.3 pounds per second; the turbine efficiency varied from 0.86 to 0,88. However, the peak turbine efficiency of 0.88 occurred over a smaller range of the performance map than at the higher Reynolds number index. Turbine performance at the lowest t-urbine- inlet Reynolds n-umber indices investigated (0.13 to 0.19) is shown in figure 10(c). The much larger range of this turbine performance map resulted from the combina- tion of closed-stator data at a compressor-inlet Reynolds number index of 0.2 (lower right portion of map) and open-stator data at a compressor- inlet Reynolds number index of 0.08 (top left portion of map). The tur- bine efficiency lines must be considered only approximate because of the relatively large Reynolds number index variation for this low Reynolds number index range; but it is evident that the peak efficiency shifted to a region of lower corrected work and pressiire ratio than at the higher Reynolds number indices. The minimum turbine efficiency encountered during the investigation was about 0.82 and occurred during closed-stator operation at a Reynolds number index of 0.20. The trend of the constant corrected speed lines to lower corrected gas flows at the low turbine pressure ratios indicates that the turbine nozzles were unchoked when operating with the variable stators in the closed position. When the turbine nozzles were choked, the corrected tiirbine gas flow was still about 28.3 pounds per second, the same as at the higher turbine Reynolds number indices. The effect of Reynolds number on turbine performance is summarized in figiu-e 11. At the conditions where Reynolds number effect could be isolated, that is, at constant values of corrected turbine speed and pressure ratio, the turbine efficiency was not affected by Reynolds num- ber down to a Reynolds number index of about 0.4, but dropped off about 0.02 as the Reynolds number index was reduced to 0.15. There was no apparent Reynolds number effect on corrected turbine gas flow over the range of Reynolds number indices investigated. Exhaust system . - The exhaust -system data are presented to allow calculation of over-all engine performance from pumping characteristics which are based on turbine -out let pressure. Tailpipe total-pressure loss data are shown in figure 12 as a function of the turbine gas-flow param- eter w ^-/rr/Pc-, which is a function of the turbine-outlet Mach number. With the variable stators in the open position, the tailpipe total- pressure loss increased from 4.5 percent of the turbine -out let total pressure at a gas-flow parameter of 1.02 up to about 13 percent at 1.47. At the static sea-level military condition, the gas-flow parameter is 1.26 and the total-pressure loss ratio is about 0.07. The turbine 10 NACA RM E58C12 gas-flow parameter could be lowered to 0.65 with the stators fully- closed, at which point the total-pressure loss ratio was about 0.04. Data obtained with the largest exhaust-nozzle area resulted in total- pressure loss ratios as high as 0.35 (not shown on figure), indicating that choked flow existed somewhere near the turbine instead of at the exhaust-nozzle throat. The exhaust nozzle became unchoked at turbine gas-flow parameters somewhere between 1.47 and 1.53. The velocity coefficient of the primary exhaust nozzle with the secondary nozzle removed is shown in figure 13 as a function of nozzle pressure ratio. When the exhaust nozzle was unchoked (nozzle pressure ratio P0/P9 "*■ 0'5)^ data scatter made the values unreliable. When the exhaust nozzle was choked, however, the data fell about a mean value of about 0.985. Engine Performance Several aspects of the over-all engine performance are discussed in this section. Typical effects of variable stator position on net thrust and specific fuel consumption are presented at a specific flight condi- tion and exhaust-nozzle area. Engine pumping characteristics with open stators are presented at selected values of Reynolds number index and the general effect of Reynolds nuniber on the pumping characteristics is summarized. Wet thrust and specific fuel consumption are presented for a flight Mach number of 0.9 at rated engine conditions over a range of altitudes above the tropopause to illustrate over-all performance losses resulting from decreased Reynolds number index. Some typical effects of variable stator position on over-all engine performance are shown in figure 14 for one specific flight condition and exhaust-nozzle area. Net thrust fell off rapidly in approximately linear fashion when the stators were closed. This is, of course, the expected trend on the basis of the corresponding airflow and compressor pressure ratio reductions (figs. 7(a) and (b)). Although the thrust dropped rapidly, the specific fuel consumption increased only slightly with closure of the stators to about' the mid position (18°). However, the specific fuel consumption increased rapidly as the stators were closed further. This nonlinear variation of specific fuel consumption with stator position is a result of the similar variation in compressor effi- ciency with stator position shown in figure 7(c). Inasmuch as the vari- able stators are generally scheduled to be open except for certain engine transient operations at reduced speeds, the following presentation of generalized steady-state engine performance is confined to the open stator position. NACA RM E58C12 11 Pumping characteristic maps, which consist of the variation of en- gine pressure ratio with corrected engine speed with lines of constant engine temperature ratio and corrected airflow, are shown in figure 15 for Reynolds number indices of 0.6, 0.2, 0.12, and 0.08. The peaks of the lines of constant engine temperature ratio show the regions of maxi- mum combined compressor and turbine efficiencies. The slope of the lines of constant corrected airflow at low corrected engine speeds re- flects the reduction in corrected airflow with increasing pressure ratio as discussed in the compressor performance section. Over-all engine performance may be determined for choked exhaust-nozzle operation at any flight condition corresponding to a Reynolds number index greater than 0.08 by use of the pumping characteristic maps and several auxiliary curves (figs. 6, 8, 12, 13, 18, and 19). A sample calculation of engine performance using this method is presented in appendix C. The general trend of engine pressure ratio and corrected airflow with Reynolds number index is shown in figure 16 for several corrected engine-speed and temperature-ratio conditions. Curves similar to these can be constructed from the pumping maps for calculating engine per- formance at other engine conditions and can be interpolated for inter- mediate values of Reynolds number index. The reduction in net thrust and increase in specific fuel consump- tion resulting from Reynolds number effects on the engine components are shown in figure 17 for the rated engine speed and limiting temperature condition over a range of altitudes from about 35,400 feet (tropopause) to 72,000 feet at a flight Mach number of 0.90. This corresponds to a range of compressor-inlet Reynolds number indices from 0.46 to 0.08. Increasing the altitude over this range reduced the corrected net thrust by 14 percent, 6 percent of which was due to reduced airflow. The lower engine pressure ratio resulting from reductions in the compressor and turbine efficiencies accounted for about 6 percent of the thrust loss, and about 2 percent was due to the increased tailpipe pressure loss brought about by the above effects on the turbine-outlet Mach number. The specific fuel consumption was increased about 16 percent as altitude was increased over this range, 6 percent of which can be charged to com- bustion efficiency, 8 percent to the compressor and turbine efficiencies, and 2 percent to the higher tailpipe pressure loss. SUMMARY OF RESULTS The results of performance tests on the XJ79-GE-1 turbojet engine and its components are summarized as follows: 1. At rated corrected engine speed (7460 rpm) and design compressor pressure ratio, the corrected airflow was 159 pounds per second and the compressor efficiency was 0.784 at a Reynolds number index of 0.6. At 12 MCA RM E58C12 this Reynolds number Index, peak compressor efficiency was between 0.80 and 0.81 and occurred at a corrected engine speed of approximately 6900 rpm. Compressor performance was not appreciably affected by Reynolds number at Reynolds number index values greater than approximately 0.4. Lowering the Reynolds number index from 0.4 to 0.08 reduced the com- pressor efficiency 0.05 and lowered the corrected airflow approximately 7 percent at the highest corrected speed (8000 rpm) at which comparisons could be made. Increasing the engine inlet temperature from 416° to 700° R at a constant Reynolds number index resulted in approximately a 3 percent higher corrected airflow for a given corrected engine speed and compressor pressure ratio. 2. Varying the variable stators from the open to closed position (0° to 35 ), resulted in reductions in corrected airflow and compressor pressure ratio on the order of 60 percent. Peak compressor efficiency was lowered only 0.02 by closing the stators halfway, but decreased rapidly when the stators were closed further. Net thrust fell off rapidly when the stators were closed, but the specific fuel consumption remained relatively low until the stators were closed more than halfway. The planned schedule of variable stator position as a function of cor- rected engine speed apparently safely bypassed the low- speed stall region and passed through the regions of peak compressor efficiency. 3. Combustion efficiency varied from 0.97 to 0.90 over the range of engine and flight conditions investigated with the variable stators in the open position. The combustor total-pressure loss ratio varied from 0.05 to 0.07 over the range of combustor temperature ratios investigated. 4. Turbine efficiency varied only from 0.88 to 0.86 for open-stator operation throughout the investigation. The minimum turbine efficiency encountered was about 0.82 and occurred during closed-stator operation at a Reynolds number index of 0.20. Turbine efficiency was not affected by Reynolds number down to a turbine-inlet Reynolds number index of about 0.4, but dropped off about 0.02 as the turbine-inlet Reynolds number index was reduced to about 0.15. There was no apparent Reynolds number effect on corrected turbine gas flow over the range of Reynolds number indices investigated. 5. An increase in altitude from the tropopause to 72,000 feet at a flight Mach number of 0.9 (Reynolds number index reduction from 0.46 to 0.08) resulted in a 14 percent reduction in net thrust and an increase in specific fuel consumption of 16 percent in comparison with the values that would be obtained assuming no losses in component performance with increasing altitude. Lewis Flight Propulsion Laboratory National Advisory Committee for Aeronautics Cleveland, Ohio, March 20, 1958 NACA RM E58C12 13 APPENDIX A SYMBOLS The following symbols are used in this report: A area^ sq ft B balance force from thrust capsule, lb C coefficient F thrust, lb g acceleration due to gravity, 32.17 ft/sec H enthalpy, Btu/lb M Mach number N engine speed, rpm P total pressure, Ib/sq ft abs p static pressure, Ib/sq ft abs R gas constant, ft-lb/ (lb) (°R) T total temperature, °R t static temperature, °R V velocity, ft/sec w flow rate, lb/sec or Ib/hr JC_ X correction factor. 1.4 V 2 / r 1.4 'r^) 1.4 - 1 X ratio of specific heats 8 ratio of total pressure to NACA standard sea-level static pressure 14 NACA RM E58C12 6/
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