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Home , Intercooler, Overall pressure ratio, Piston

E AMEICA SOCIEY O MECAICA EGIEES 92-GT-405 4 E. 4 S., ew Yok, .Y. 00

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C. OGES Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1992/78941/V002T02A028/2401669/v002t02a028-92-gt-405.pdf by guest on 23 September 2021 Sundstrand Power Systems San Diego, CA

ASAC NOMENCLATURE High altitude long endurance unmanned aircraft impose KFT Altitude Thousands Feet unique contraints on candidate propulsion systems and HP types. , rotary and have been proposed for such special applications. Of prime importance is the HIPIT High Pressure Intercooled Turbine requirement for maximum (minimum specific Mn Flight Mach Number fuel consumption) with minimum rejection. Engine weight, although secondary to fuel economy, must be evaluated Mls Inducer Mach Number when comparing various engine candidates. Weight can be Specific Speed (Dimensionless) minimized by either high degrees of turbocharging with the Ns piston and rotary engines, or by the high power density Exponent capabilities of the gas turbine. pps Airflow The design characteristics and features of a conceptual high SFC Specific Fuel Consumption pressure ratio intercooled are discussed. The intended application would be for long endurance aircraft flying TIT Turbine Inlet Temperature °F at an altitude of 60,000 ft.(18,300 m). It is estimated that such a Weight lbm turboprop would be capable of thermal efficiencies exceeding 40% with current state—of—the—art component efficiency levels EI Effectiveness and an overall pressure ratio of 66.0. Projected E Regenerator Effectiveness Power (at altitude) to weight ratio is comparable to that of competitive piston and rotary engines. AP/P Cycle Pressure Drop Ratio

rntd t th Intrntnl G rbn nd Arnn Cnr nd Exptn Cln, Grn n 4, 2 1.0 INTRODUCTION This paper describes the extension of the HIPIT concept studies to very high altitude long endurance aircraft applications. The study of propulsion systems for very high altitude It discusses engine design point performance optimization, part prolonged endurance aircraft is addressed in References 1.2 and load performance, and preliminary engine design configurations. 3. In summary. these reNrences indicate:

• The major engine design constraint is low specific fuel 2.0 GAS TURBINE CYCLE OPTIONS consumption (SFC). Gas turbine engine thermal efficiencies greater than 40% • SFC's below 0.30 lb/hp-hr (thermal efficiency 44%) are (SFC 0.33 lb/hp.hr) can be obtained by the incorporation of desired. compressor intercooling and/or exhaust heat regeneration.

• Engine design is optimized at the high altitude (60 KFT or Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1992/78941/V002T02A028/2401669/v002t02a028-92-gt-405.pdf by guest on 23 September 2021 The simple cycle (non-heat exchanged) gas turbine has higher) loiter condition, with ambient air temperatures as low dominated in aircraft applications, since the additional weight as -70°F. and volume of the heat transfer surfaces becomes excessive for • Engine weight can be reduced without compromising SFC. typical aircraft operational profiles. Since higher thermal efficiencies are paramount to the operational success of Both cooled spark ignition piston and gas turbine prolonged endurance aircraft, comparative cycle computations engine propulsion systems have been successfully demonstrated were analyzed for intercooled, intercooled and regenerative, and in high altitude aircraft. This is confirmed by the Boeing regenerative engine configurations. Figure 2 shows the results "Compass Cope" and "Condor" aircraft research programs. and the cycle computations for operation above the tropopause, Rotary Wankel stratified change internal combustion engines are and from which the following is concluded: also being studied. The major attribute of both the piston and Wankel engines is essentially flat rated output power with • All three engine configurations are estimated to provide altitude made possible by high pressure turbocharging and/or SFC's in the 0.3 lb/hp.hr realm. turbocompounding. In contrast, gas turbine turboprop output power decreases rapidly with altitude (air density) as typified by • Higher pressure ratio improve the straight intercooled cycle the power lapse rate conditions shown in Figure 1 for constant performance. speed and turbine inlet temperature. Further power decrements • Lower pressure ratios improve the regenerative and occur as a function of operation at extremely low Reynolds regenerative intercooled cycle efficiency. numbers. When constrained to operate at constant speed and turbine inlet temperature with a Reynolds exponent of 0.09, • Turbine inlet temperature has less influence on straight Figure 1 indicates a power lapse rate less than 0.1, accompanied intercooled engine cycle efficiency than the regenerative by increasing SEC. This would imply that a typical turboprop . rated at 4000 hp at sea level would deliver less than 400 hp at Reference 5 indicated the feasibility of a very efficient gas altitudes approaching 60 KFT turbine cycle incorporating both intercooling and regenerative are, however, capable of delivering very high heat exchange, with moderate pressure ratios for a twin engine specific power hp/lb engine weight and dominate in aircraft light aircraft application. As yet, however, the propulsion applications above 500 hp. has found limited usage in aircraft applications. It presents more mechanical complexity than the lower temperature intercooler. The results of a trade study of various propulsion systems for Fundamental problems with high temperature heat exchangers high altitude, long endurance aircraft are described in Reference are creep, and eventual leakage causing engine performance 1. An advanced intercooled and regenerated gas turbine engine deterioration. The high exhaust volume flows at very high concept was estimated to be capable of delivering SFC's below altitudes also tend to either increase the gas side pressure drop or 0.30. Further investigation of the concept was, however, enlarge the gas side flowpath. recommended to assess geometry and weight of the intercooler and regenerator. Table 1 lists the advantages and disadvantages of all three options. Although the intercooled regenerative cycle Independent studies have been conducted by the author on a incorporates the best of both approaches, it results in the most high pressure intercooled turbine (HIPIT) described in complex flowpath. and also incorporates the worst of both Reference 4. approaches - the regenerator.

2 TABLE 1. COMPARISON OF ENGINE CONFIGURATIONS

TYPE ADVANTAGES DISADVANTAGES

Intercooled • No high temperature heat • Requires very high pressure ratios transfer surfaces • Reduced leakage potential • High number of stages • Small, responsive high pressure • Intercooling spool • Interco°ler heat rejection Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1992/78941/V002T02A028/2401669/v002t02a028-92-gt-405.pdf by guest on 23 September 2021 • High specific air power

Regenerative • Low pressure ratio, reduced • Low specific power number of stages • Requires high effectiveness with very large regenerator • High leakage • Poor transient response • Variable geometry for best part load fuel economy • Fouling

Regenerative • Lower stage number than • Complicated ducting to Intercooled intercooled both intercooler and regenerator • All regenerative issues including fouling

In effect, the choice lies between large hot surface power (hp/w pps) and specific weight (hp/lb), plus aircraft regeneration with simple (single—stage) turbomachinery, lower endurance trends. cold surface intercooling with multistage turbomachinery, or a combination of both. The utilization of very high cycle overall pressure ratios, resulting in lower final exhaust temperatures, has suggested the The use of ultra—high pressure ratios also alleviates the possibility of exhaust reheat. Exhaust reheat would occur prior problem of maintaining high combuster efficiency at high to the final work production expansion stage for power boosting altitudes. Figure 1 shows that the combustor loading parameter purposes during intermittent emergency mode operation. increases up to five times that at sea level. A pressure at 60 K The concept is based on a Brayton cycle with very ratio of at least 30:1 would be required to maintain optimum high overall pressure ratio and compressor intercooling. This loading for a given combustor volume. combination produces high thermal efficiency, high specific power and excellent part—load fuel economy.

.0 HIP1T CONCEPT Extensive baseline studies of the HIPIT concept have been conducted (Ref 5). They verify that the approach is uniquely Recent applications of small radial (centrifugal) capable of meeting future challenging thermal efficiency and turbines have demonstrated their capability to deliver very directives. high overall pressure ratios when triple spooled with . For example, with effective intercooling and a A conceptual HIPIT engine schematic arrangement is shown stage pressure ratio of approximately 4.0, an overall compressor in Figure 4 and comprises a triple spool arrangement with a pressure ratio of 64 has been demonstrated. The attainment of separate free power turbine. Intercoolers are positioned after such high overall pressure ratios in combination with high cycle the first low pressure (LP) compressor, and after the second temperatures offers the possibility of improved Brayton cycle intermediate pressure (IP) compressor. Reheat may be used thermal effectiveness beyond those values shown in Figure 3. between the LP turbine and power turbine for power boosting. These are competitive with those of the Diesel cycle. The The smaller high pressure spool is offset aft behind the potential advantages of the high pressure intercooled turbine intercoolers, and provides an acceptable frontal area approach (HIPIT) offer an alternative to the lower pressure ratio for aircraft applications. intercooled and heat exchanged cycle. A HIPIT cycle analysis code was prepared to parametrically examine the tradeoffs In—line design of all turbine stages (and reheat combustor), between specific fuel consumption SFC (lb/hp—hr), specific with the exception of the high pressure spool, provides a clean

"back end" flowpath. A forward-mounted intercooler is upon compressor characteristics and matching. The triple spool depicted to possibly take advantage of in-flight ram. arrangement results in each stage operating along independent operating lines permitting operation near the compressor peak efficiency islands. 4.0 CYCLE PERFORMANCE Since stage pressure ratios exceeding approximately 6.0 Design optimization normally begins with a design normally result in the encroachment of stress limits for both the requirement for a gas turbine of given power output, specific fuel compressor and turbine, the HIPIT cycle analysis was initially consumption, and specified weight and size. Within these influenced to stage pressure ratios of 3.0, 4.0 and 5.0 with turbine confines, it is customary to select an optimum combination from inlet temperatures of 2460°R, 2860°R and 3260°R. Equal work the following cycle variables: split was selected for the HIPIT compressor stages. Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1992/78941/V002T02A028/2401669/v002t02a028-92-gt-405.pdf by guest on 23 September 2021 • Turbine Inlet Temperature (TIT) Weight assessment was computed in the cycle analysis and - to Ambient Temperature Ratio included the following components: • Compressor and Turbine Efficiencies • Compressor and turbine assemblies • Pressure Ratio • Intercoolers • Combustor Efficiency and Pressure Drop • Combustor primary and reheat • Interco°ler and Heat Exchanger Effectiveness • Power turbine • Parasitic Losses (Gearbox, Bearings, Ducts) • Gearbox

Comprehensive performance analysis for small gas turbines • Accessories using single-stage radial compressors and turbines has been Intercooler effectiveness versus size and weight trends follow developed (Reference 6 and 7) and extended to include the the general form: constraints of turbine rotor stress rupture life.

Current methods of predicting the peak component Size and Weight oz 1-E (1) efficiencies demand lengthy computation procedures and extensive input, including complete turbo-machinery geometry Since aircraft applications are normally of shorter duration description and performance requirements. Such prediction than ground or marine applications, and are also influenced by methods are too inflexible for use in a cycle optimization inlet air momentum drag considerations, optimum values of procedure. To obtain a practical procedure, it is necessary to effectiveness for aircraft are normally on the order of 0.65 to define component efficiencies in terms of a reduced number of 0.75. parameters without sacrificing accuracy. The major parameters which influence component efficiency are rotational speed, Cycle analysis results are shown in Figure 7 for the range of operating clearances, and state-of-the-art level. stage pressure ratios and T.I.T.'s, with intercooler effectiveness of Et or 0.85, 0.75, and 0.65, and include proportionate allowance The influence of rotational speed and compressibility can be for fan driven intercoolers removing up to 20 percent of the assessed from specific speed charts of typical single-stage radial compression generated heat. Specific fuel consumptions are in flow compressors and turbines shown in Figures 5 and 6. the 0.3 to 0.40 lb/hp.hr range, and compete with those of the Component state-of-the-art efficiency levels are depicted, piston and rotary engines. based upon the defined limitations. They can be digitized for inclusion into cycle analysis routines, together with the effects of Total propulsion system power-to-weight ratios up to physical size as influenced by Reynolds number effects. 0.75hp/lb are indicated for the lower intercooler's effectiveness Compressor specific speeds of 1.0, 0.85 and 0.62 were chosen for and higher turbine inlet temperatures, with slightly increased the HIPIT, low pressure. medium pressure (MP) and high SFC. pressure (HP) stages. Higher single-stage compressor pressure ratios are realized by increasing rotor tangential velocities, but Before selecting a design point cycle condition from the SFC also incur higher Mach numbers at the rotor entry and.exit. The and specific weight trends, it is worthwhile to study the influence aerodynamic problems associated with diffusion at these of both of these important parameters on aircraft endurance. conditions are being resolved to attain efficient high pressure ratio compressor operation. These problems involve careful Endurance Considerations selection of the blading solidity, thickness chord ratios, nose An abbreviated form of the Breguet equation for expressing radius, hub and shroud contours, appropriate rotor and diffusion the relationship between endurance and initial take off (wi) to ratios, and strict control of the design dimensions. final landing weight (W2) for a given aircraft is:

At higher stage pressure ratios, compressor surge and turbine Endurance is inversely proportional to SFC, with lowest SFC matching do not always allow operation at peak compressor yielding maximum endurance. The fuel fraction for long efficiency. Therefore, design point compressor efficiency may be endurance aircraft normally exceeds 0.5W1, compared to the one or two percentage points below peak efficiency, depending propulsion system weight which may be as low as 0.1Wi.

4 combustor. The shaft-driven cooling fan is still depicted, but 0.5 could equally be either electrically driven, or possibly ram Endurancecv Constant W1 -1 I (2) cooled. SFC W2 Propulsion system weight represents a second order influence on The high pressure spool is a critical component, and is in endurance, but similar to all aircraft components, must be effect a miniature small gas turbine in its own right. It is similar to scrutinized to maximize payload and range. the U.S. Army sponsored T100 "MPSPU" engine (Reference 8) shown in Figure 10. This high pressure spool is envisioned as The examination of long endurance aircraft system weights driving an /, and engine accessories. Its and performance from Reference 1 indicates that the influence small size would minimize battery sizing and would provide very of engine SEC and specific power (HP/lb) on aircraft endurance rapid response to power transients. can be approximated by: Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1992/78941/V002T02A028/2401669/v002t02a028-92-gt-405.pdf by guest on 23 September 2021 Engine design point performance conditions for this Constant ri configuration are listed in the first column of Table 2. The overall Endurancea HP (3) SFC klbJ pressure ratio selected at the 60 KFT altitude design condition was 66.0 with a T.I.T. of 2000°F, and a net SFC of 0.30 lb/hp-hr. where the exponent n is on the order of This assumes current state-of-the-art attainable component 0.15-0.25. efficiency levels shown with typical intercooler and combustor A relative assessment of the influence of both SFC and pressure drops. specific power on endurance, using equation 3, is shown in Figure 8, with pressure ratio, intercooler effectiveness, and T.I.T. Operating the engine in "flat-rated" mode, i.e. constant as parameters. power with altitude, would back the spools to the sea level condition listed in the second column of Table 2, with an Examination of Figure 8 shows the optimum intercooler overall pressure reduced to 12.3 and a T.I.T. of only 1400°F In all effectiveness for maximum endurance is in the range of probability, such an engine would not be designed to function at 0.65-0.75. Endurance increases with increasing pressure ratio full throttle sea level conditions in order to minimize casing and to a lesser degree with increasing T.I.T. The optimum T.I.T. is weights and gearbox power limitations. Excess power would, 2400°F. Although elevated T.I.T.'s are feasible, with internal however, be available for increased take-off weight depending cooling or ceramic materials for the HP spool turbine, it would upon capability. be prudent to limit the loiter T.I.T.'s to the lower level of 2000-2200°F. Additionally, some slight temperature margin is In this respect, it should be noted that the previous engine required for power transients. weight estimates were based upon component weight correlations from typical existing turboprop engine designs, and may therefore be somewhat pessimistic for this unique Design Characteristics and Features application. For example, the LP and MP compressor impellers The previous cycle optimization studies were important in could be made from aluminum rather than titanium, even the last identifying candidate cycle aerothermodynamic design turbine stage might be made from a high temperature titanium conditions. However, additional factors require consideration alloy. in the overall engine design formulation process. These include: The part load performance was conducted matching all three • Engine mechanical design features. compression and four expansion stages on the basis of flow • Requirements for aircraft systems power supplies. continuity and work balance. The results of the analysis are shown in Figure 11 where normalized specific fuel consumption • Unique structural aspects with maximum pressure ratio at and overall cycle pressure ratio are plotted versus normalized extremely low ambient pressure. power ratio. The part load SFC curve is quite flat with an SFC increase of only 10 percent at one-third power. Optional use of "Part load" operation at sea level and lower altitudes. • "top end" reheat would boost rated power 35 percent with an • lntercooler number, size, and packaging versus effectiveness increase in SEC. of 20 percent. tradeoffs. Analysis of the individual compressor stage matchings A conceptual mechanical design arrangement for a 400 hp showed the operating lines in Figure 12 would provide positive pusher propeller installation is shown in Figure 9. The surge margin. This would be acheived without the requirement arrangement is similar to that of Figure 3, with the addition of the for variable geometry or interstage bleed with the possible propeller reduction gearbox and exclusion of the reheat exception of the LP spool at very low powers.

AE 2. CAIAE I ESIG OIS

AMIE 60 K S.. (AKEO Compressor Stage Pressure Ratios 4.8, 3.94, 3.78 2.5, 2.4, 2.3 Compressor Stage Efficiency 0.79 0.78 Intercooler Effectiveness 0.75 0.75 Interco°ler Pressure Drop(s) 4 4 Overall Compressor Pressure Ratio 66.1 12.3

Combustor Pressure Drop(s) 6 6 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1992/78941/V002T02A028/2401669/v002t02a028-92-gt-405.pdf by guest on 23 September 2021 Combustor Efficiency(s) 98 99 Turbine Inlet Temperature(s) °F 2000 1400 Turbine Stage Efficiency 86 86 Mechanical Efficiency 95 95 Leakage 1.0 1.0 Airflow pps 1.65 4.9 Gross Output hp 432 437 Cooler Power hp 32 37 Net Output hp 400 400 SFC lb/hp.hr 0.3 0.51 LP Rotational Speed krpm 26.1 18.9 LP Compressor Tip Diameter in 13.3 MP Rotational Speed krpm 40.7 29.4 MP Compressor Tip Diameter in 8.6 HP Rotational Speed krpm 56.5 41.0 HP Compressor Tip Diameter in 6.2 Power Turbine Pressure Ratio 11.2 2.4 Estimated Weight lb 805

Conclusions Cycle optimization considered both the engine thermodynamic and aircraft performance constraints in the Gas turbine engine conceptual design studies were tradeoffs involved between engine specific fuel consumption and conducted to assess the feasibility of an advanced technology engine specific weight. intercooled turbine engine for high altitude long endurance propulsion applications. Engine weight estimates were based upon conventionally designed components and are expected to be conservative for this unique application. Engine weight, however, was shown to Cycle performances of regenerated, regenerated plus be of lesser importance than thermodynamic efficiency. It intercooled. and intercooled gas turbines were computed. All requires attention, as indeed do most weight elements of the long three types. when fully optimized, were capable of thermal endurance aircraft. Propulsion system power (at altitude) to efficiencies exceeding 40%, and are therefore viable candidates weight ratios of between 0.5 and 0.8 were projected. These are for high altitude long endurance applications. The specific competitive with piston and options. merits of each type were presented. In effect, the choice lies between large hot surface regeneration with simple (single stage) It was shown that the optimum intercooler effectiveness for turbomachinery, lower cold surface intercooling with multistage maximum endurance is in the range of 0.70 to 0.75. It was also turbomachinery, or combinations of both. A preference was demonstrated that T.I.T.'s higher than 2400°F provide minor made here for the straight intercooled configuration with endurance extension, apart from material life and manufacturing moderate intercooler effectiveness to minimize heat rejection cost considerations. The heat rejection issue for the intercoolers and cooler volume/weight. A preliminary conceptual design was is the major concern for this proposed gas turbine propulsion generated with an overall pressure ratio of 66.0 achieved with system. It was partially addressed by estimating and deducting three intercooled single—stage centrifugal compressors. A bonus cooling fan losses from the gross output power. Continued of high pressure ratio is improved combustor performance at studies to refine optimization of the intercooler effectiveness as a extremely high altitudes. function of type, weight, volume, and the possible addition of

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ram cooling or heat recovery with a small Rankine bottoming REFERENCES cycle are suggested. 1. Kowalski, E.J., Baullinger, N.C., Kolden, J, 1991. "Propulsion System Endurance for an Unmanned High The specific application and its unique requirements result in Altitude Long Endurance RPM" ASME 91-GT-409. a specialized propulsion system with very limited commercial 2. Wilkinson, R.E., Benway, R.B., 1991. "Liquid Cooled viability coupled with extensive design and development costs. Turbocharged Propulsion System for Hale Application." Conversely the highly turbocharged piston or rotary engine ASME 91-G-399. appear to offer some degree of commonality with existing product lines and hardware. This development philosophy may 3. Gallington, R.W. 1991. "Propulsion Requirements for also be applied to "turbocharging and intercooling" an existing High Altitude Long Endurance Flight." ASME 91-GT-393. small core gas turbine as mentioned previously. It could also Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1992/78941/V002T02A028/2401669/v002t02a028-92-gt-405.pdf by guest on 23 September 2021 apply to adapting an existing larger turboprop with intercoolers 4. Rodgers, C, 1988. "High Pressure Interco°led Turbine with a high pressure offset (non-concentric) core. The offset Engine Concept". SAE 881204. core concept resolves the mechanical difficulties involved with multiple concentric shafts at the ultra high pressure ratios 5. Mock, E.A., Caldwell, R.T., Boyd, K.E., 1984. "Regenerative Interco°led Thrbine (RITE) Study". envisioned . AIAA-84-1267. 6. Rodgers, C, 1991. "The Efficiencies of Single Stage for Aircraft Applications". ACKNOWLEDGMENT ASME 91-GT-77.

The author wishes to acknowledge the efforts of Robert 7. Rodgers, C, 1990. -Review of Mixed Flow and Radial Geiser in programming the cycle optimization procedure; Nancy Turbine Options", AIAA 90-2414. Roussakis, Diana Prince and Mary Greenawalt in preparation of 8. Rodgers, C., Napier, J.C., Thompson, R.G., 1989. the manuscript; and Sundstrand Power Systems for publication "Development Test Status T-100 Multipurpose Small authority. Power Unit", ASME 89-GT-117.

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