RESEARCH OF A HIGH THRUST-TO-WEIGHT RATIO SMALL TURBOFAN ENGINE

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1 THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS 345 E. 47th St., New York, N.Y GT-289 The Society shall not be responsible for statements or opinions actvancedin papers or discussion at meetings of the Society or of its Divisions or Sections, or printed in its publications. Discussion is printed only if the paper is published in an ASME Journal. Authorization to photocopy material for Internal or personal use under circumstance not falling within the fair usenmvisions of the Copyright Act is granted by ASME to libraries and other users registered with the Copyright Clearance Center (CCC)Transactional Reporting Service provided that the base fee of $.3 per page is paid directly to Ihe CCC, 27 Congress Street, Salem MA 197. Requests for special peimission or bulk reproduction should be addressed to the ASME Tedrecal Publishing Department CopyrigtttO 1997 by ASME All Rights Reserved Printed in U.S.A RESEARCH OF A HIGH THRUST-TO-WEIGHT RATIO SMALL TURBOFAN ENGINE Masahiro Akagi iii Masashi Shinomiya Third Research Center, Technical Research & Development Institute Japan Defense Agency Junichi Sakaki Shunji Sugai Aero-Engine & Space Operations Ishikawajima-Harima Heavy Industries Co., Ltd. Abstract, The 3rd Research Center of the Technical Research and Development Institute (TRDI) of Japan Defense Agency (JDA) and Ishikawajima-Harima Heavy Industries Co., Ltd. (IHI) developed and tested the demonstrator of a high thrust-to-weight ratio small turbofan engine with an afterburner called "XF3-4, the purpose of which is to establish engine technologies for the future supersonic aircraft for JDA. The development program started in 1981 and the first engine test was carried out in All the engine tests planned completed in March 1995 successfully. This paper reports the design, development and test results of the XF3-4 engine above. Nomenclature AB:Afterburner AFT: Afterburner Fuel Pump AMU: Afterburner Management Unit A8: Exhaust Nozzle Atea CDP : Canpressix Diathaige Posure CID: ComputerFluidDynamics EGT : Exhaust Gas Temperature FADEC: Full Authority Digital Elects:de Cato! MMU: Main Management Unit Nf : Fan Speed Ng: Compressor Speed PLA : Power Lever Angle Psi Fan Inlet Stat Pressure Ttl : Fan Inlet Tanperature 1TI' : Turbine Inlet Temperature 12. Canpressor Inlet Temperature VSV : Variable Stator Vane VSVA: VSV Angle Wfab: AB Fuel Flow Wf: Main13umer Fuel Flow Introduction The general view of the XF3-4 is shown in Fig. 1. The target of.thrust-to-weight ratio was set at 7:1, which was the highest level in the world for the small turbofan engine of 34kN class rated thrust as shown in Fig.2. It shows that the smaller the engine thrust becomes, the smaller the thrust-to-weight ratio becomes. It is difficult for smaller engines to achieve high thrust-to-weight ratio, because they have some structural restrictions and aerodynamic disad- Fig.I XF3-4 ENGINE vantages as follows. First, small engines have some structural and manufacturing limitations of weight reduction. The accessories' size cannot be reduced proportionally to the engine size. The minimum metal thickness of casting parts for large engines needs to be also applied to small engines to keep good casting quality. The small engines are relatively heavy because those materials are therefore relatively thick in comparison with large engines. Secondly, small engines have a lot of disadvantages on aerodynamic performance, such as the Reynolds number effects on component efficiencies, boundary layer effects on compressor rear stages and relative air leakage through the Presented at the International Gas Turbine & Aeroengine Congress & Exhibition Downloaded From: on 1/1/219 Terms of Use: Orlando, Florida June 2-June 5,1997

2 THRUST-TO-WEIGHT RATIO XF3-4 MAXIMUM LEVEL OF THE MILITARY ENGINES IN SERVICE ' THRUST (4N) fl' 81 ' '85 '87 89 '9 '91 '92 '93 '94 Component Development Test Compressor Combustor HP Turbine Afterburner Full Scale Engine Design Itsla Engine Test Fig.3 DEVELOPMENT PROGRAM Fig.2 RELATIONSHIP BETWEEN THRUST AND THRUST-TO-WEIGHT RATIO clearance between casings and blade tips, compared with large engines. The demonstrator engine incorporates such advance technologies as three-dimensional CFD aerodynamic design for the compressor and turbine -airfoils, high efficiency cooling design for the turbine blades and nozzles, high heat release rate of the combustor, new materials like single crystal for the turbine blades and nozzles, powder metallurgy for the turbine disk, ceramic bearings for the AGB, a high reliable FADEC system and so on in order to achieve the target of high performance. The thrust-to-weight ratio of 7:1 was achieved and the engine performance was confirmed as designed. The advanced technologies proven in this program will be applied to the future JDA engine development program. Development Program The development program of the XF3-4 is shown in Fig.3. lb confirm the design method and the performance of the components, the following component tests were conducted. Compressor : The research of a high performance compressor was initiated in 1981 and the full five-stage compressor had been tested on a rig since 1987 and the some stages were modified by the 3rd Research Center and THI from 199 to 1991 to satisfy the requirement of the adiabatic efficiency and the surge margin. Combustor : In order to obtain the data of the ignition and blow off limits, basic tests had been conducted with rectangular combustor models since From 1987 to 1991, sector combustor models were tested and annular-type combustors were supplied to the performance test by the 3rd Research Center and THI. HP Turbine : Since 1985, HP turbines for the m of 17K had been researched. From 199 to 1991 heat-cycle tests were conducted by the 3rd Research Center and Till. In 199, disks made of powder metallurgy were manufactured. From 199 to 1991, the disks were supplied to cyclic spin tests at the 3rd Research Center and Afterburner: lb define the afterburner stability limits, tests with afterburner sector models were begun in Then the sector afterburner models were tested from 1985 to 1988, and some afterburners were installed into the F3-THI-3 and had been tested since 1986 at THI and since 1988 by the 3rd Research Center. Engine Test : The engine tests were successfully conducted at the 3rd Research Center from 1992 to 1994 and the total engine run time was accumulated to about 25 hours. Design Features General Arrangement : The major engine characteristics of design are listed in Table 1 and the general layout is shown in Fig.4. The major features of its architecture are as follows: -a two-stage wide cord fan a five-stage compressor with variable inlet guide vanes and variable 1st stage stator vanes 2 Downloaded From: on 1/1/219 Terms of Use:

3 Table 1 Thrust(Dry Max) 21.4kN(2,184 (Max AB) 34.21cN(3,49kg SFC(Dry Max) 21mgNs(.76kgfihAtgt) (Max AB) 61mgiNs(2.15kgf/h/kg Thrust-to.Weight Ratio 7 Overall Pressure Ratio 14 Bypass Ratio.9 Turbine Inlet Temperature 1673K AS Exit Temperature 1973K Maximum Engine Diameter 66mm Overall Length 2,729mm Fig.5 COMPRESSOR STATOR VANE sor airfoil is shown in Fig.5. The threedimensional CFD analyses were also applied to the HP and the LP turbine for high load controlled vortex airfoil design. Fig.4 THE ICF3-4 GENERAL LAYOUT a high load combustor with air blast type fuel nozzles a high loading single-stage high-pressure (HP) turbine made of a single crystal with advanced cooling structures and a single-stage low pressure (LP) turbine a chute type mixer and a spray bar type fuel injector of an afterburner a light weight convergent variable exhaust nozzle a full authority digital electronic control (FADEC) with a hydromechanical backup control 3D Aerodynamic Design : Three dimensional computational fluid dynamics (CFD) analyses were applied to design the controlled diffusion on the compressor blades and vanes in order to satisfy the target of the high pressure rise per stage and the high efficiency. The end bent was also applied to minimized the vortex loss near the wall. The pressure rise per stage of 1.4 is the highest level among the engines currently inservice in the world. The sample of the compres- High Load Combustor : The high load and high performance combustor was newly developed for the XF3-4 engine to reduce the engine size and weight. The heat release rate is J/m 3/s/Pa, which is the highest level in the world as shown in Fig.6. In order to realize such a compact combustor, the air flow in the dump diffuser and the combustor was analyzed and trimmed for an optimum flow distribution between primary and secondary air using CFD and rig test results. Ceramic coating was applied to the machined ring liner which is shown in Fig.7. HEAT RELEASE RATE (J/s/m '/Pa) X HEAT RELEASE RATE 195 Wt' Hu HIGHEST LEVEL OF THE MILITARY ENGINES IN SERVICE con ( V Ps k S M3 Pa W1 : FUEL FLOW Hu : LOWER CALORIFIC VALUE V : LINER VOLUME P3 : COMBUSTOR INLET PRESSURE J XF YEAR Fig.6 TREND COMBUSTOR PERFORMANCE 3 Downloaded From: on 1/1/219 Terms of Use:

4 SHOWER HEAD RETURN FLOW COOLING \ IMPINGEMENT COOLING DIFFUSER HOLE A FILM COOLING IMPINGEMEN Ik\ COOLING CONVECTION COOLING Fig.? COMBUSTION LINER 1 1/4 TURBULENCE PROMOTER VANE tt COOUNG AIR BLADE Advanced Cooling Design of HPT : The turbine inlet temperature (Ill) of the XF3-4 is 17K level, which is the highest level for the small engines of less than 51EN thrust as shown in Fig.8. The vanes and blades were made of single crystal casting. The return flow cooling structure was incorporated into the blades and the full-converge film cooling was applied to the vanes as shown in Fig.9. These cooling design achieved 17K of TIT without thermal barrier coating. High Load Afterburner : The temperature rise in the afterburner is 12K and the fuel-to-air ratio is 1:17. The afterburner module consists of Fig.9 COOLING TECHNOLOGY OF HPT BLADE a chute mixer, three segments of spray bars, that is, Local, Fan and Core, and a flame holder with a pilot burner. Fig.1 shows the spray bar arrangement looking forward from the exhaust nozzle. The chute mixer attains good mixing between the core flow and the bypass flow under the minimum pressure loss Fig.11 shows the results of CFD analyses of the flow mixing at the chute mixer. The shape of the chute was selected to give the good flow conditions at the flame holder for afterburning. (K) 22 2 tg IF= - a cm 12 7 P XF3-4 Ct.." -. o..., --' ".,6 HIGHEST LEVEL OF THE MILITARY ENGINES IN SERVICE g THRUST 4.14) Fig.8 RELATIONSHIP BETWEEN THRUST AND TURBINE INLET TEMPERATURE Fig.1 AFTERBURNER (LOOKING FORWARD) 4 Downloaded From: on 1/1/219 Terms of Use:

5 A ANALYTICAL MODEL(LOOKING FORWARD) FLOW SECTION B-B. 4-- FLOW Fig.12 ROLLERS OF CERAMICS BEARING level of reliability and safety. Fig.13 shows the FADEC, which is designed to be compact and light-weight incorporating advance technologies SECTION A-A FLOW PATTERN Fig.11 CFD ANALYSIS OF CHUTE MIXER AGB Ceramic Bearing : All the bearings in the accessory gear box are ceramic bearings shown in Fig.12. The balls and rollers of the bearings are made of ceramics, silicon nitride. It enables to reduce the lubricant oil flow. Advanced FADEC Control System: The engine control system consists of a FADEC, a MMU, a AMU, two VSV actuators and four exhaust nozzle actuators. The FADEC controls the main combustor fuel flow, the VSV angle of the high pressure compressor, the exhaust nozzle area (AS) and the AB fuel flow. The MMU contains the back-up hydromechanical control of a dashpot type control of the accelldecel fuel flow and the VSV angle, which holds the minimum level of the engine control function for flyhome capability even in the case of the FADEC failure. The control system realizes the high - Fig.13 FADEC like surface mount on printed circuit boards. The exhaust nozzle area actuators are operated by the fuel pressurized by the main fuel pump and the APP to eliminate an additional hydraulic system dedicated only to the nozzle control. There were three AB fuel metering valves for a Local, a Core, a Fan AB fuel segment in the MMU which were controlled independently in order to hold the AB burning in the most suitable condition over the full range of the AB operation. Engine Test Fig.14 shows the XF3-4 installed on the engine test bed. The engine tests were conducted successfully and the total engine run time was 5 Downloaded From: on 1/1/219 Terms of Use:

6 ) 4 8 ti 3 DRY MAX DESIGN POINT MAX AB DESIGN POINT Fig.14 XF'3-4 INSTALLED ON THE TEST BED accumulated to about 25 hours including the AS burning run time of 1 hours. The total number of cycle of IDLE to DRY MAX. amounted to about 5 with no significant damage. The following are the remarkable results obtained in the engine test above. 2 1 DRY 1 2 Fn(kN) : MEASURED 3 4 Fig.16 STEADY STATE PERFORMANCE (Fn vs SFC) Steady State Performance : The engine steady state performance data of Fn vs EGT and SFC vs Fn are shown in Fig.15 and 16 respectively. The measurement data are in good agreement with the design and prediction. 25-2% - /Th 2 45 E-1 PREDICTION ESIGN POINT % : MEASURED WfF/VVfAB 5 MEASURED EGT(K) Fig.15 STEADY STATE PERFORMANCE OF DRY ENGINE(EGT vs Fn) Optimization of AB Zoning Schedule: It is necessary for the AB control to optimize the fuel distribution between the core segment and the fan segment. Fig.17 is the data obtained from 6 Fig.17 OPTIMUM AB FUEL DISTRIBUTION the optimization test of the AB zoning schedule. It shows that there is an optimum distribution around the WfF/WfAB of.4, that is, 4% of the total AS fuel is supplied to the Fan segment. The temperature distribution at the exhaust nozzle exit was also measure& Fig.18 shows the data obtained under the optimum fuel zoning schedule above at the maximum AB rating. Fig. 19 illustrate measuring the exhaust gas temperature with the water-cooled traverse probe. Downloaded From: on 1/1/219 Terms of Use:

7 O O O ' O : PLAdee Fig.2 AS SCHEDULE OVER THE DRY RANGE Fig.I8 TEMPERATURE DISTRIBUTION AT EXIT OF EXHAUST NOZZLE UNDER MAX AB CONDITION 5 1 TIME(sec) Fig.19 SCENE OF MEASURING CONDITION QUANTITY AT THE EXIT OF EXHAUST NOZZLE UNDER MAX AB CONDITION Transient Performance : The exhaust nozzle is modulated at Idle according to the control schedule given in Fig.2. This control logic enables to increase Idle speed (Ng) holding the Idle thrust less than 5% of Max Dry as specified in MIL-E-57D. In result, the engine transient operation from Idle to Dry Max was achieved in 4.2 seconds as shown in Fig.21. The acceleration was recorded smooth and stable without any engine surge nor any hazardous EGT overshoot etc. This control logic shortened the acceleration Fig.2I TRANSIENT PERFORMANCE IDLE TO DRY MAX(X-T, PLA, Fn, Ng, AS) time by.7 second from the conventional control logic. In order to reduced in the transient time from Idle to Max AB, a part throttle AB logic, that is, AB operation before reaching Dry Max, was introduced. The engine transient operation from Idle to Max AB was achieved in 5.5 seconds without any significant problems as shown in Fig.22. These transient performance satisfied the MIL-E.57D's requirements. Electrical Power Failure Test : A dashpot control system was newly developed as the hydromechanical back up control. It controls the main engine fuel flow and the compressor VSV angle only, using a dashpot mechanism scheduled by Psi and Ttl. The purpose of the control 7 Downloaded From: on 1/1/219 Terms of Use:

8 ISO 95916F pia " Fa 5 1 TrotInt) Fig.22 TRANSIENT PERFORMANCE IDLE TO MAX AS (X-T, PLA, Fn) is to provide 8% of the Dry Max thrust for flyhome capability over the specified engine flight envelope in the case of FADEC failure. Conclusion 15 The XF3-4 engine was developed and tested successfully. It achieved the target of thrust-toweight ratio of 7:1, incorporating the advanced technologies like 3D aerodynamic design, a high oo load combustor, high efficient cooling design, 2 new materials and so on. All the planned test were carried out successfully. It was confirmed so through these tests that the engine steady state and transient performance satisfied the design target and MIL-E-57D's requirements. The advance technologies proven in this program will be applied to the future JDA engine development program. Reference I. I.Kashikawa, M.Akagi, S.Yashima and M.Ikeyama, "Research on a High Thrust-to- Weight Ratio Small Turbofan Engine", AIAA , July 1995 Aero-Engine & Space Operations Ishikawajima-Harima Heavy Industries Co., Ltd. Downloaded From: on 1/1/219 Terms of Use: