Installation Requirements for Small Gas Turbine Engines

THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS 345 E. 47 St, New York, N.Y. 10017 85-IGT-151 The Society shall not be responsible for statements or opinions advanced in papers or in 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. Papers are available from ASME for fifteen months after the meeting. Printed in USA. Copyright 1985 by ASME Installation Requirements for Small Gas Turbine Engines T. F. PIERCY C. E. HEATHCO Allison Gas Tubine Division General Motors Corporation Indianapolis, IN 46206 ABSTRACT Allison currently has over 20,000 Model 250 small gas turbine engines operating throughout the world in a variety of turboprop and turboshaft installations. The experience base has contributed to the understanding and development of small engine installation requirements. Some of the critical interface systems will be reviewed to highlight some of the small engine considerations of engine installation design. NOMENCLATURE rpm Btu g Hz in./sec RPR TDC shp MGT %Ap INTRODUCTION revolutions per minute British thermal units acceleration force frequency inches per second ram pressure recovery top dead center shaft horsepower measured gas temperature percent inlet pressure loss SMALL ENGINE CONFIGURATION The Model 250 engine design was initiated to provide a high specific power gas turbine engine for both fixed- and rotary-winged aircraft. Some of the major criteria established during the design phase included the following: 300 hp class significant power growth potential reliable/durable powerplant light weight compact layout field maintainable adaptable configuration installation flexibility The configuration that best satisfied these requirements is the center drive, center exhaust layout with compressor and turbine modules located on either side of the accessory gearbox. With this arrangement, the air is drawn into the compressor, compressed, and discharged through a scroll into two external ducts that convey the air to a single can type combustor. During combustion the flow changes direction and discharges into the turbine located between the combustor and gearbox. After expanding through the gas producer and power turbines, the air exits through the exhaust collector, as shown in Figure 1. The gearbox provides the primary structure for mounting the engine in the aircraft. Installation requirements for small gas turbine aircraft engines have evolved from the individual requirements from various administrative, technical, and commercial sources. These requirements are intended to ensure that the integration of the engine into the aircraft provides a reliable, safe product that will satisfy customer expectations. While many of these requirements are generic to all classes of gas turbine engines, certain ones are specifically established for the unique features of the small engine class. Some of the major engine to airframe system integration areas will be reviewed here to illustrate how the requirements were established and how they are applied to the Model 250 engine, a widely used small gas turbine engine built by Allison Gas Turbine Division of General Motors. The basic engine configuration (which will be a point of reference here), the application to specific aircraft types, and the interaction between the engine and the aircraft are subjects for examination to illustrate the contribution that each has made in establishment of detailed installation requirements. Fig. 1. Model 250-C28B turboshaft engine. Presented at the 1985 Beijing International Gas Turbine Symposium and Exposition Beijing, People's Republic of China September 1-7, 1985

A turboshaft engine version, a Model 250-C20B, shown in Figure 2, features a top outlet exhaust system and a 6016 rpm power takeoff drive speed for connection to a helicopter main rotor reduction gearbox system. By applying the adaptable configuration requirement, the same basic engine is adapted for turboprop use, as illustrated in the Model 250-B17C engine of Figure 3. The components are reoriented to a downward exhaust arrangement and a propeller reduction gearbox is added to reduce the 6016 rpm output to the 2030 rpm that is a typical small propeller speed. TURBINE ASSEMBLY COMBUSTION ASSEMBLY Fig. 4. Bell OH58D helicopter. ACCESSORY GEARBOX ASSEMBLY Fig. 2. Model 250-C2OB engine. PROPELLER REDUCTION GEARBOX Air Intake Systems Fig. 5. Government aircraft factories Nomad N-22. COMPRESSOR POWER TURBINE COMBUSTION AND SECTION ACCESSORY GEARBOX Fig. 3. Model 250-B17C engine. The Model 250 engine has grown from the original 250 hp YT63 engine (Model 250-B1) to include four series of commercial (Model 250) engines and three series of military (T63 and T703) engines with power ratings up to 650 hp. More than 20,000 engines of this family have been produced for use in 27 turboshaft and 26 turboprop applications. Examples of the light aircraft types using this small engine are shown in Figures 4 and 5. AIRCRAFT SYSTEM INTEGRATION The Model 250 engine applications experience has shown that certain areas are critical to completing a successful small gas turbine installation. The specific factors addressed here are the following: air intake systems lubrication system interface loads Air intake system design and evaluation are significant to ensuring the satisfactory performance and function of gas turbine engines. Because the air intake pressure and temperature conditions directly affect available engine power, internal engine stress levels, and proper engine function, intake system design is perhaps the most critical area of engine installation. Many of the following items are considerations from the engine viewpoint in the intake system design: inlet pressure losses inlet pressure distortion inlet temperature increase inlet temperature distortion ram pressure recovery foreign object ingestion protection ice formation aircraft type and configuration The relative importance of these factors is dictated by the type of inlet system used on the aircraft. In practice two general categories of inlet systems are used with small turboprop or turboshaft engines. The plenum inlet is normally applied in rotary-winged installations to maintain relatively constant low velocity inlet conditions with widely varying aircraft attitudes, flight direction, and airspeeds. The helicopter requirements also include inlet particle protection because of the frequent operation close to the ground with the accompanying potential of foreign object ingestion. On the other hand fixed-wing aircraft inlets place more importance on the performance of inlet design. Here the attainment of minimum pressure losses and maximum ram pressure recovery are the primary goals because of the higher airspeeds desired. In either inlet system type the particular design is assessed for temperature and pressure conditions that exist at the engine compressor inlet with survey equipment such as is shown in Figure 6. The assess- 2

ment includes determination of both the average levels and spatial variation of these parameters. Analysis of the extent and criticality of the variations that exist at the compressor inlet are done to compare to the manufacturer's distortion limits that are established during engine development. The concern for minimizing these losses may be illustrated by the data shown in Figure 8. Zero loss and 5% power loss performance curves are shown. To overcome the 5% losses and to restore the power to the original no loss condition, the engine would need to be operated at approximately 50 F (28 C) higher turbine measured gas temperature. In a cruise power condition this turbine temperature increase would reduce turbine component life by approximately one half. 450 0. 400 5% loss 0 350 0. O fa 300 250 Fig. 6. Compressor inlet survey instrumentation. Typically the manufacturer has limits addressing the following items with corresponding concerns: inlet pressure distortion compressor blade or vane dynamic loading compressor surge (stall) margin installed performance power stability/response inlet temperature distortion installed performance compressor surge (stall) margin power stability/response Inlet condition evaluation is particularly important because the small engine may be sensitive to minor flow surface discontinuities that would be normal and acceptable in larger engine installations. For example, a surface variance of perhaps 0.050 to 0.100 in. may represent a 5 to 10% change of flow annulus height in a small engine and potentially excite a vibratory response in the compressor. In the worst case a premature engine failure could result. Experience has shown that early coordination and joint engine/airframe design review can minimize this type of problem. 200 1100 1200 1300 1400 1500 Measured gas temperature F Fig. 8. Model 250-C20B performance. To illustrate the interaction of inlet configuration on pressure distortion, pressure drop, and ram recovery, three typical inlet systems are compared in Figure 9. The plenum shown in Figure 9a has excellent static pressure distribution and a ram recovery of 94%. This system also has 5 F temperature rise for a total installed static power loss of over 8%. This same plenum with a protective screen is shown in Figure 9b; loss due to the screen is approximately 1.5%. The duct in Figure 9c is a direct inlet with turning just aft of a propeller. Due to supercharging from the propeller, a static pressure rise is shown in a segment of the inlet. Total ram recovery of this system is nearly 99% with negligible temperature rise. The manufacturer usually further provides recommendations as to the levels of average inlet pressure loss or inlet temperature rise that are consistent with engine installed performance considerations. Sensitivity of these parameters is shown for the Model 250 engine in Figure 7., 31.0 30.5 30.0 - Prof c 200 knots 98.8% RPR 2.7% PP c static 99.3% RPR 4%AP O 3 C 2 2 flc z 1 2 3 4 5 8 7 8 9 10 INLET DEPRESSION Inches 120 INLET TEMPERATURE RISE. f --- EXHAUST BACKPRESSURE Inches 020 Fig. 7. Engine performance sensitivity. g 29.5 29.0 a 28.5 - TDC 90 180 CIRCUMFERENTIAL LOCATION 270 380 --a 150 knots 94.2% R PR 3%AP :1 '7% RPR 1% AP.0 static 95.2% RPR 1% AP Fig. 9. Compressor inlet pressure variation. The other major factors, foreign object ingestion protection and inlet icing protection, are major design considerations because failure to incorporate such provisions may result in either mechanical damage to the engine or loss of some operational capability with the engine. One approach employed for these is to prevent entry of foreign particles or ice by inertial separation of the heavier particles from the inlet air. However, turning of the air during separation adds additional inlet flow system losses with resulting power losses. An alternate method is to provide a barrier such as a screen to prevent particle entry but with large enough openings to minimize performance losses. Alternate air entry 3

paths are also required with barrier types to always ensure adequate inlet air in the event of primary system blockage. The compressor inlet areas on small engines contain anti-icing provisions to prevent accretion of ice on stagnation areas. The aircraft inlet system should incorporate similar provisions to prevent inlet blockage from icing with potential compressor damage and inlet distortion. Both electrical resistance heaters and engine bleed air have been employed for anti-icing inlets. In some installations a screen is located ahead of the inlet to prevent larger ice particle entry. The screen is shaped so that air may be drawn around the sides when ice forms on the screen. An ice tolerant inlet system is illustrated in Figure 10. In another aspect of lubrication interface, the aircraft system must provide the capability to remove entrained air from the scavenge return oil prior to returning the oil to the engine inlet. Because the scavenge oil pump capacities by problem definition must exceed that of the supply pumps and because the air from the labyrinth air/oil seals used in engines enters the oil system, there is always a large quantity of air entrained in the scavenge oil. A major function of the oil tank is to remove this air quickly and efficiently. Features such as baffles and trays are incorporated into the aircraft oil tank to increase oil dwell time and achieve air removal. Failure to address this issue will result in the ability to provide adequate oil flow and pressure to the engine. The problem manifests itself in loss of pump efficiency and is evidenced by reduced oil flows. Allison is currently adding a new small engine installation requirement to the lubrication system based on both operational experience and inhouse development tests. The addition of an external 27 micron or finer scavenge return oil filter to the lubrication system shows improvements to engine bearing and gear life, increased oil change intervals, and reduced internal carbon formation over that resulting with the normal 80 micron engine oil filter. The substantial improvement in carbon formation rates is shown in Figure 11..050 - Lubrication System Fig. 10. Screened ice tolerant inlet. In small gas turbine engines such as the Model 250, the lubrication system integration into the aircraft can be a constant source of problems to the user or can ensure a trouble-free, satisfactory, reliable product, depending upon the approach and attention given this area. The basic performance of the oil pump pressure and scavenge (return) elements is dependent on the inlet oil temperature, inlet oil pressure, air/oil content, and scavenge oil back pressure that are influenced by the components used in the design of the installation..040.030.020 STANDARD OIL SYSTEM-N1/4 A major area of concern to the engine is that of basic system cooling. Small engines featuring relatively high oil flow and heat rejection rates, which are inherent with the compact physical size of this engine class, require careful definition and assessment of this portion of the lubrication system. The small engine configuration, with the close proximity of the high temperature gas path components to the rotor bearing cavities, requires that the engine designer use a major portion of the total lubrication flow for internal cooling of the bearing cavity areas. This design solution places an additional requirement on the user of small engines by requiring a cooling system size that approaches that required for engines several times the small engine power class. This is illustrated by a comparison of two Allison turboprop engines in Table 1, which shows the oil system parameters that the installer must consider in sizing the cooling system. Table 1. Comparison of Allison turboprop engines. Power rating Heat rejection Engine model hp Btu min Oil flow lb/min 501-D13 3900 3335 154 501-D13 2000 2715 154 250-B17D 420 1440 32 250-B17D 200 1120 32 As a further illustration of the scope of the oil cooling problem, a small engine with a 4 gal/min engine oil flow rate and a 1.5-gal oil supply tank will completely recirculate this oil system every 15 sec. At the same time the temperature of the oil must be reduced by approximately 100 F (55 C) prior to reentering the engine..010 ENDURANCE TIME-HOURS Fig. 11. Carbon accumulation rates. Establishment of this filter as an installation requirement versus incorporation as part of the basic engine package is a good example of the installation coordination required in small engines, particularly when the design and configuration are established. Because of the established installation of the many aircraft using the engine and the size of the filter package, providing a kit that can be used by all users of the engine is practically impossible. In reality, it is a better commercial solution to define the technical installation criteria and allow the users to achieve the solution that suits their particular installation and application. Interface Loads The engine manufacturer recognizes that aircraft loads in many forms will be transmitted to the engine. The task then is to identify the load types, design to accept them, and to place limits where necessary to protect the engine and ensure the durability and operating safety standpoints. The loads addressed in the installation design may be grouped into several categories such as the following: flight maneuver and aerodynamic loads vibratory input loads 4

Flight Maneuver and Aerodynamic Loads Flight maneuver loads are generally static or low cycle vibratory loads that result from either aerodynamic input loading or actual aircraft maneuvers. These g loads are the result of actual aircraft flight and are intended to be applied through the engine mounting structure. The general procedure for the engine builder is to design and substantiate the product during development to achieve the requirements specified by the customer and airworthiness agencies and to evaluate the usage of the product to ensure compliance. The loads generally are designed to ultimate failure conditions (i.e., crash) and do not usually impose any problem or burden for routine conditions of use. Verification of these loads is a combination of test measurements during operation over the flight envelope and analysis of these data for other conditions such as projected aircraft mission to demonstrate compliance. Flight evaluation is necessary to verify analytical assumptions of load application in the airframe structure and engine mounting arrangements. Aerodynamic loads may, however, cause concern to the installation, particularly in small turboprop engines. The propeller used in these applications allows continuous steady loading to the engine from the forces created by airflow at an angle relative to the propeller disk. Since this load varies with aircraft angle of attack, airspeed, gross weight, center of gravity, etc, the magnitude of this load must be established and incorporated in the flight load analysis. Further, if any maneuver loads, such as aerobatic spins or stalls, are allowed, these loads must be defined and shown to be within structural and mechanical system limits. The engine manufacturer may offer several mounting options to provide installation flexibility of the product. An example of the possible mount combinations for a new Model 250 derivative engine is shown in Figure 12. Depending on the anticipated load requirements, the aircraft structure proximity, the maintainability requirements, or other installation criteria, the installer may select which of the five configurations best addresses his installation. The potential benefits include reduced installation weight and improved access to the engine for installation/ removal. Vibration Loads Small gas turbine engines feature rotational speeds of more than 50,000 rpm, several times that of large gas turbines. With these high rotational speeds the engine designer is concerned that external aircraft-generated vibratory loads may be input to the engine and cause undesirable engine structure deflections, which, in turn, cause misalignment loads to critical engine components. The engine designer defines an installed vibration level that has an incremental level increase available for aircraft-generated vibrations. Included are propeller or main rotor rotational vibrations, accessory vibrations, vibratory aerodynamic loadings, and engine high frequency response to these inputs. To assess the effect of aircraft vibration on engine vibration, an installed vibration survey is a necessary. By placing vibration transducers at key engine locations, the engine carcass vibration modes may be assessed for amplitude and frequency response. Those modes that respond may then be addressed with isolation techniques to reduce or translate the vibration forcing function. The objective of this survey is to ensure that unexpected loading is not applied to the engine. Figure 13 shows vibration limits for a Model 250 engine. Note that over the normal range of engine frequencies, approximately 100-2000 Hz, an overall installed vibration level of 1.5 in./sec. average velocity is allowed. Acceptable test cell vibration levels range from 0.6 to 0.9 in./ sec. Low frequency vibrations (below engine sources), typically below 40 Hz, are defined in terms of 1.5 g to accept the large displacements that are experienced here. 5.0 01 15 40 100 1000 2000 Frequency, Hertz Fig. 13. Vibration limits. BOTTOM MOUNT FRONT GEARBOX PAD ALLOWABLE MOUNTING SYSTEMS 2 SIDE WITH BOTTOM 2 SIDE WITH TOP 2 SIDE WITH REAR TURBINE FRONT WITH BOTTOM FRONT WITH REAR TURBINE REAR TURBINE MOUNT SIDE MOUNT (2 PLACES) Fig. 12. Model 250-C2OP turboshaft mounting options. These limits have evolved from engine development experience and operational lessons learned. The current limits represent an acceptable compromise between limiting the engine high frequency (low deflection) levels and addressing the airframe low frequency (high deflection) vibrations. Power absorber system torsional vibration is another primary installation vibration concern. Main rotors and propellers are coupled to the engine through gearboxes and lightweight shafting that often have low frequency natural torsional resonant frequencies similar to engine control system response frequencies. Cyclic torque variations that may exist in these conditions require thorough evaluation during the installation testing. These torque pulses, which are sensed by the engine governor and affected by control system response rates, power absorber hysteresis, and drive system inertias, may create a divergent torque response in the shafting as the governor tries to maintain a desired engine speed. The governor must therefore be tolerant to low frequency inputs to prevent this condition. Control system stability can be achieved by establishing the engine control system response rate at a different frequency than that of the power absorber. In the Model 250 engine, pneumatic control system 5

volumes are tuned by selecting appropriate accumulator sizes or other response attenuators. While system math models and prior flight experience are helpful in initially selecting the control volumes, the effects of altitude load application rates and aircraft maneuvers must be verified in flight tests. Torsional vibration evaluation has in some installations been expanded to include some higher frequency conditions. In this case the reverse concern exists where the engine excited a response in the aircraft shafting. Because of the potential of excessive bearing, gear, shafting, and structure loading, a dynamic shafting torsional survey has been added to some Model 250 engine installation tests. SUMMARY The small gas turbine engine is a complex product that has successfully been applied as a light aircraft powerplant. Part of the reason for the success lies in the attention given to the installation design requirements that have been established by the engine manufacturer, end product users, airworthiness agencies, aircraft builders, and others. The contributions of these parties have established unique tests and standards that have been applied to this engine class, have resulted in an airworthy product, and have ensured future growth of the small gas turbine engine. It may be observed from the small engine installation history that problems will occur but that a sound engineering design and test program will ensure that a satisfactory product will result. 6