Advanced General Aviation Turbine Engine (GATE) study

SECTION 1.0 t SUMMARY i A 5-task study program was conducted to develop a market scenario, to evalu- ate the benefits of gas turbine power in general aviation aircraft in the late 1980's, and to outline the technologies requisite to meeting the market needs. The study spanned fixed and rotary wing aircraft in the 726 kg to

TT

3629 kg (1600 to 8000 lb.) take-off gross weight (TOGW) markets.

Task I showed that potential U.S. national engine sales of 31,500 per year L i t (95 percent fixed gyring, 5 rercent helicopter) could exist for defined catego- ries of the market. A primary constraint is the ability to produce engines with a sales price approaching current reciprocating engines. Fuel conserva- tion, installation, safety, comfort, and environmental improvements would accrue if the market could be c-^eated via reduced engine prices.

These features were used as input i Task II to evaluate & spectrum of engine ► configurations; optimum aircraft-engine types and payoffs were developed for each aircraft category, based on a (limited) life cycle cost criterion. It was determined advanced component design that a combination of features, a { 4 approach, and high rate pro- low cost Manufacturing and materials technology duction would produce engine designs to meet the price, performance, and dur- ability objectives.

Task III provided the derivation of a common core engine applicable across , A 9:1 pres- the spectrum, and the benefits and tradeoffs associated with it.

sure ratio single stage compressor, combined with a novel reverse flow vapor- izer plate combustor, a 1504 degree K (2250 degree F) uncooled radial turbine rotor, and a multi-purpose reduction gearbox constitute the major elements of the common core.

Commonality of these parts across a 197-422 k1a (265-565 hp) power range for turboprops and turbishafts was demonstrated.

In Task IV, a 5-year component development and plan was constructed for It was engine demonstration focusing on the requisite advanced technologies.

a plan; concluded that a successful engine family could result from such identify and reduce the risk inherent NASA's investment in the program would in the advanced aerodynamic, materials, and structural technologies.

i N A SECTION 2.0 r INTRODUCTION i Efforts are underway to improve all types of General Aviation engines (reci- procating, rotating, diesel and gas turbines). The drivers are the expanding S market, the need for energy conservation, and the demand for more stringent environmental controls.

Turbine power has been accepted (Figure 1) in larger fixed and rotary wing aircraft (above approximately 9.9 kN (2200 lb thrust) and 418 kW (560 hp) because of its benefits to flight speed, payload and aircraft gross weight, and to general passenger comfort and safety. Time Between Overhauls (TBO) intervals are currently significantly higher than competing reciprocating engines.

It was, therefore, appropriate to examine the requirements and technologies for all sizes of advanced General Aviation turbine engines that might be expected to come into service in the late 1980's. These engines include fixed (connected) and free-shaft turbines, as well as smaller size turbofans.

The specific objective of the General Aviation Turbine Engine (GATE) study was to define the requirements for small engine advanced technology suitable for General Aviation service in the 1987-1988 timeframe. Small engines are defined as being in the 112-746 kW (150-1000 hp) range; 1/2 to 3/4 of the effort was directed to engines in the 112-447 kW (150-600 hp) range. For turbofans, emphasis was on 6.7 kN (1500 lbs) thrust or less. The study included fixed and rotary wing aircraft applications, a component technology assessment effort, and a core demonstrator plan to provide the technology base to enter engineering development in 1988.

The study evaluated the opportunities of turbine power in General Aviation aircraft and generated information necessary for the Government to formulate the most effective technology program for smaller sized turbine engines.

The study was divided into four tasks. Task I focused on a market analysis to identify applications, mission profiles, and environmental requirements.

Task II encompassed trade-off studies to identify the optimum engine aircraft technology requirements. A common core concept was evaluated in Task III to assess its benefits and penalties for application across the wide range of propulsion requirements. A conceptual "optimum" core engine design results from this task, and a program plan was developed in Task IV for the follow-on component technology and core engine demonstration.

I SECTION 3.0 TASK 1: MARKET ANALYSIS Task I concentrated on projecting fixed and rotary wing markets through 1988 by functional segments using extrapolated and postulated reciprocating, rotary, and turbine engine characteristics. The projection was iterated to determine the effect of various engine technological. possibilities (weight,, size, performance, cost) and environmental noise and emissions regulations on market quantity and product distribution. From the market scenario, domains of engine superiority were identified and the distribution of applicable gas turbines (and their requirements or features) were extracted. Figure 2 des- cribes the overall approach taken to project the market potential of small gas turbines in the late 1980's time frame.

3.1 Preliminary Data Base Because the market analysis plan was being performed prior to the availabil- ity of detailed engine or aircraft specifications, it was necessary to stipu- late market categories and engine characteristics on the basis of current history and previous studies in the General. Aviation field extrapolated to the 1988 time frame. Initiation of the task was, therefore, based primarily on executive or experiential judgment. Senior Beech and Teledyne CAE person- nel utilize d their experience and prior studies of a similar nature to define bands of: a) engine potential. - weight, SFC, cost and bulk and, b) aircraft features - desirable mission payoffs, comfort features, safety, and marketable price.

Relative to engine potential, prior Teledyne CAE engine design studies were assessed, and preliminary performance analysis were accomplished, using bands of potential pressure ratio, turbine inlet temperature capability, and compo- nent efficiencies to define the SFC potential. Special consideration was given to small engine design limitations. [weight estimates were based on current production engines and the potential improvement of new materials of higher strength, cost, weight, and scaling information were assembled.

With respect to the aircraft assessed their corporate case, Beech personnel history and the limitations implicit in the types of aircraft. In the rotor- craft case, Bell, personnel identified new markets of opportunity - i.e., new products filling a need anticipated for the 1988 time frame.

tti 3.1.1 Aircraft Categories Five aircraft categories were selected, as summarized in Table I. These 3$ R(^ categories were separated by their primary mission requirements, rather than on an arbitrary gross weight basis.

^a The helicopter category shown on Table I is described in Section 3.3.

The -first category consisted of a single engine, two-passenger aircraft designed principally for flight training activities. It is a simple training aircraft and will be flown by inexperienced pilots from a fixed base opera- tion. Therefore, it should not have excessive altitude and speed capabili- ties or be complex; it should be very conservative as to aerodynamic sophist- ication.

The second category was composed of single engine aircraft with a four-pas- ^ senger capability. Category III included single engine aircraft with a four to six passenger configuration. These aircraft are most often identified as the high performance end of the single engine aircraft group. Category IV was composed of twin engine aircraft. At the end of the spectrum, the Cate- gory V aircraft was primarily a corporate airplane and could accommodate the best potential advanced aircraft and engine technology. (Category V, histor- icaI. trend data, was the only one to include twin engine turboprop aircraft.)

Later reassessment of the aircraft missions and marketable features elicited the opinion that single engine, pressurized, high performance Category III aircraft were beginning to appear in several product lines. This observation was included in the study by dividing Category III aircraft into pressurized (IIIF) and unpressurized (IIIU) types.

An agricultural aircraft was added, alth-ough it was outside the current Beech market spectrum. In each case, with payload fixed by passengers and baggage, it was necessary to complete the mission definition with cruise altitude, velocity and range assumptions. upon reviewing Task II, the initial Task I projected increases in these parameters proved to be optimistic.

For clarity of presentation, the final categories, their current price, and their missions (i.e., determined following Task II iteration) are summarized on Table I. The differences between the original judgment and final analyti- cal category definitions represent the iterations undertaken to balance Task I "desired" features and Task II "affordable" features. The final result was a 7.7 to 15.4 m/s (15-30 ktas) increase of cruise velocity, and a 10-25 per- cent increase of range, depending on category, defined for 1988 GATE-powered aircraft.

3.1.2 un ine Performance As noted previously, it was necessary to utilize executive judgment to postu- late engine technology levels which could be achieved in 1985. Assumptions were therefore made on the basis of preliminary cycle calculations, using extrapolated component technology levels for simple, single-spool turboprops and two-spool turbofans.

A single spool (connected shaft) turboprop was used in all fixed wing studies because it was judged to offer the minimum production price potential..

The mission assumptions and preliminary turboprop engine cycle definitions are summarized in fable II and Table III. Table II shows that cruise SFC varies Tess than 3 percent with a 55 degree K (100 degree F) change of tur- bine inlet temperature; however, Table III shows that required component size (flow) changes 10--16 percent. They also show that the engines will be flat- rated, i.e., to match reciprocating engine characteristics, takeoff will occur at power levels considerably below maximum temperature capability.

Sufficient information is presented to relate cruise requirements to sea is defined for subsequent air- level static capabilities, i.e., "lapse rate" craft-engine sizing.

assumed, and fan For the turbofan, similar core engine characteristics were pressure ratios were optimized at each flight condition in a preliminary ana- lysi.s. was predicated in order to Initially, a 6:1 bypass ratio turbofan Table IV sum- sizing and scaling information.

obtain the earliest possible marizes the turbofan data.

Again, the engines are seen to be flat-rated because of the low flight speed (compared and the high takeoff to higher performance turbofan aircraft) thrust available from the 6:1 bypass ratio cycle.

Accessory power requirements were defined as 1.5 kW (2 hp) for Categories I and II, 2.2 kW (3 hp) for Categories III and IV, and 6 kW (8 hp) per engine for Category V. A requirement was also defined for 0.038 kg/S (5 lb/min) bleed airflow for Categories III and IV, and 0.045 kg/S (6 lb/min) for Cate- gory V. power requirements were considered in sizing the engines for These each category, i.e., the cruise power was increased by the accessory require- ment. Bleed air was assumed to be provided by a separate gear driven Com- discus=ed in Sec- pressor. The effect of this approach versus bleed air is tion 4.2.

Bell for integra- This range of performance data was submitted to Beech and tion with the engine sizing and cost analysis to define aircraft-engine com- binations.

Cost, Weight and Scaling Data 3.1.3 EnFine for 1985 engine technology as an Engine scaling and cost data were developed for a turboprop:` input to the marketing analyses. These data were developed Task I turboprop is a single shaft and a turbofan engine. The GATE baseline centrifugal compressor, a' engine (Figure 3) consisting of a single stage and a reduction, gear. The reverse-flow annular combustor, a radial turbine, scaling data turboprop installation is shown in Figure 4, and the baseline geared turbofan are are shown in Table V. Similar data for a single shaft shown in Figures 5, 6, and Table VI.

The basic scaling equation is expressed as follows: N = N,, (PWR/PWR,, )x where N = dimension or weight x = exponent defined for each dimension or weight PWR = Kil-owatts (horsepower) or thrust as applicable The dimensions and These scaling data were generated from softline sketches.

power, Caere reduced to exponential functions, weights, as a function of thereby and consistent scaling data over the power providing continuous range. The 116 KW turboprop is beyond the scaling range, thus a single tabu- lation for this size is shown in Table V.

comparison with current technology engines.

The engine was by cost estimated a t The Task x baseline turboprop is compared with a 1976 technology hypothetical production turboprop in Figure 7: the significant difference in the schematic, sections is the reduction in the number of compressor and turbine stages.

The engine component cost, relative to the hypothetical production turboprop engine is shown in Table VII.

Detailed component cost data be highly proprie- are generally considered to tary. The baseline component estimates shown here are based on a combination of both private and published data and do not represent a detailed analysis of a specific engine. considered both appropriate 1 These data, however, are cost estimation and sufficiently accurate to be useful in the GATE engine procedure.

All of the following comparisons are made at a fixed engine weight, irrespec- tive of engine cycle or performance. The latter factors are addressed later.

f Comparing GATE advanced technology to the hypothetical production engine, the gearbox and air inlet housing are estimated to be equal. The compressor cost is reduced from stages have been reduced from two to one; the relative 0.16 to 0.08. "vaporizer plate" design The combustor features a simplified to replace an atomizing nozzle, and the cost reduction is estimated to be 33 percent, from 0.06 to 0.04 (based on Teledyne CAE in-house studies). Threel turbine stages have been replaced by one; the cost is reduced from 0.20 to 0.07.

The cold housings (compressor section) and hot housings (combustor and turbine section) have been reduced due to the reduction in number of compres- sor and turbine stages. The accessory system cost is primarily the fuel con- trot. GATE to replace envisions using a full authority electronic control the current hydromechanical cost reduction from controls, with an estimated 0.24 to 0.12. (A&T) column has been reduced, due to The assembly and test the fewer components, from 0.06 to 0.04.' The resulting cost is 60 percent of the hypothetical production engine.

?; The third comparison features advanced fabrication technology, such as powder metal gears and squeeze and turbine rotors, and die cast aluminum housings cast compressor rotors for a further cost reduction from 0.6 0.5. The to estimated OEM (Original Equipment Manufacturer) price of the hypothetical production turboprop engine is $ 74,000 in 1977 dollars. The Task I GATE advanced technology baseline engine price would be 60 percent of this, or r $44,400.

'r When the engines are scaled to different sizes, there will be a price change as a function of size. The price relationship as a function of engine weight if is shown by the CR (cost ratio) curve in Figure 8 (based on Teledyne CAE gen- erated data for scaling a fixed configuration over a size range). The cost formula representing; the curve is shown in Figure 9. The $44,400 price pre- viously developed is for a GATE turboprop engine weighing 149.8 kg (330 lb.). j Normalizing to a CR of 1.0 and a weight of 227 kg (500 lb.), the OEM price would be $59,870. The single shaft turbofan (Figure 5) is similar to the turboprop (Figure 3); the compressor, combustor, and turbine are the same, 1 the high ratio turboprop reduction gear is replaced by a smaller, low ratio gear and a fan stage. The turbofan cost (for the same weight) is assumed to be equal to the turboprop. The turbofan and turboprop core are the same, the small high speed turbofan reduction sta ge are assumed to equal g p gear and fan sta q the price o f the heavier low speed reduction gear of the turboprop.

at the The engine price is based on current technology turboprops produced substantial market penetra- rate of approximately 500 per year. Assuming a by a factor tion of the GATE engines, these production rates would increase would accrue from the high pro- of 10 to 20, and substantial cost reductions dnction rates. as a function of production The estimated relative price rates is shown in Figure 10. The 500 per year starting point is representa- tive of current turbine production rates, and the step change at 2000 per year assumes the new fabrication technology would be implemented at this point. The slope of the curve is based on a 90 percent improvement curve for the number of units i n one the yearly production rate produced year. As increases, an additional cost reduction is expected due to increased automa- tion. This improvement curve is estimated to be 80 percent, i.e., doubling the production rate reduces the price to 80 percent of the initial value.

The final price equation is expressed as follows: C K K 1 2

(Rx b

/Ri ) a price at X units per year C K Where: C i = price at i units per year K 1 ratio of "retail" or "list." price to OEM (Original Equipment Manufacturer) price K 2 fabrication: technology factor R.rate ( units per year) corresponding to 3.

the price Cx R x rate ( units per year) corresponding to the price Cx^ b exponent based on the slope of the improve- ment curve For the dual improvement curves 90 and 80 percent) the exponent b = 0.474.

( The factor K1 has been estimated at 1.5. The value of K2 is 1.0 for Rx less than 2000 units per year and i s estimated to be 0.80 for the reduction gear and air inlet, and 0.60 for the compressor and the turbine sections for Rx greater than 2000 units per year. The composite factor for the engine is 0.833.

3.2 Market Research and Analysis - Beech: Fixed Wing Aircraft Activities in the market research task were divided into four sub-projects as shown on Figure 11: Proper identification of aircraft categories. The criterion was that o category were closely related, aircraft mission profties in each e.g., flight training in Category 1. (The categories were discussed in Section 3.1.1).

ri o Engine-aircraft price relationship analysis. The output of this task today's was a current market cost per kilowatt (horsepower) for engines, and engine price--to-aircraft price relationship in the cur- rent market.

li >i o From this, equa- An aircraft price-to-demand relationship analysis.

tions relating the current market aircraft volume to aircraft price were developed.

a.

o H i storical demand trends for the aircraft categories, extrapolated to ^I [ ] 1988.1 The output of each project is combined to define a perturbed (GATE-influ- enced) 1988 market.

3.2.1 Aircraft - Engine Price Relationship Analysis, The turbine engine initially considered for replacement in each of the defined aircraft categories provided power ratings at cruise equal to those currently in. production. The engine cost data previously developed was used to calculate engine price. utilized historical delivery Since the study volume trends spanning 14 years, it includes the effect of changing airframe tec`'.nology and new models. New models which would be developed in the 1977 and later timeframe would have these effects built into the forecast. _!

In the engine historical price analysis, only those aircraft manufactured by Cessna, Beech, or Piper were studied because of the accessibility of the engine information. engine Within each category, the individual aircraft power and price were identified, and the corresponding kilowatt cost per (horsepower) and engine percent of aircraft price were calculated. The aver- age values for each Category are shoran in Figure 12. A relationship between engine and aircraft price for each of the aircraft categories was constructed by comparing the 24 percent value for turboprop powered Category V aircraft to the 14 percent for reciprocating powered aircraft and adding this 10 per- cent increment to the historically observed percentages for the other catego- ries.

Engine price to aircraft price percentages were used to derive the maximum acceptable engine price for each aircraft price in each category, with the current size engine used in each aircraft. Combining 1) the cost per kilo- watt (horsepower) for equivalent turbine engines for each category, 2) the above engine price to aircraft price relationship, and 3) an aircraft volume z_i to price relationship, the resulting 1988 disturbed market forecast was der- ived through simple iterations of the equation: PCA = P14R x $/PIM - Aircraft Price Where: PWR = Kilowatts or Horsepower PCA = Percent Aircraft Price Th' i l h bl ^s equation was terated until t e PCA was aqua to t e maxamum accepta e h value shown in Figure 12, thereby defining the market penetration.

As noted in Section: 3.1, this modeling technique used executive judgment on the part of Beech staff, because detail aircraft and engine price analysis would not be available until Task 11. It, therefore, remained for Task II to validate the apparently simplistic assumptions that a marketable airplane would result - this later proved to be the case, as described in Section 4.5.2.

3,2.2 Aircraft Price - Demand Approach Because of overlapping price ranges in Categories I, II, and III (Table I) the three categories were grouped. The result is a more realistic price-de- mand curve that eliminates the influence of small volume airframe manufactur- ers and other abnormal perturbations in the historical delivery figures.

Within each of the categories (I-III, IV, V) the aircraft were grouped by price. In the case of categories I-III, the . aircraft were grouped into five price ranges starting at $16,000 through $91,000 -- in $15,000 increments.

The corresponding 1976 delivery volume for each of these price ranges was determined. For example, the first group consisted of those aircraft priced between $16,000 and $31,000 -- with a combined 1976 delivery volume of 5569 airplanes. These groupings were then modeled from scatter plots and replot- ted on a price -vs- demand graph as shown in Figures 13 through 15.

3.2.3 Historical Trends The output from the historical demand analysis included a 1988 undisturbed (no GATE turbine influence) market forecast, derived from a straight line extrapolation of the historical trends in delivery volumes for each category of aircraft. Typical values are shown in Figures 16 through 19. In each extrapolations were computer-modeled frog+ 5, 10, and 15 year historical case, data to evaluate the effects of national economy trends. As typical exam- ples, the figures show that the 1966-68 timeframe represented high sales volumes, whereas sales in 1970 72 (a time of recession) were low. Incorpora- - tion or exclusion of these cyclic trends resulted in large changes in the 1988 end-point of the projection. The 1988 undisturbed market projections are also shown in Figures 16 through 19. Based on prior GAM and Beech esti- mates, the 10 year values were selected as most valid for Categories IV and V. However, the growth in Categories I through III (single engine aircraft is expected to exceed the 10 year extrapolation as shown in Figure 16. The second output from the historical demand trends was a perturbed forecast of 1988 delivery volumes for each category, as influenced by the 1988 introduc- tion of a low cost turbine powerplant suitable for each of the aircraft cate- gories. This perturbed forecast is discussed further in the next section.

3.2.4 Price -Demand Analysis A simple undisturbed forecast of 1988 deliveries of Category I aircraft is 3100 units (Table VIII). It was deemed that penetration of this market by i i e ^ .s since most of the activities of turbine powered aircraft Mould be minimal, these aircraft are flight instruction. This conclusion was drawn on the ,,{ basis of the low sophistication level of the aircraft, and more so on the expectation that turbine prices in this market (power levels of 74.6 kW 3100 !,,,f hp) could not result in a sufficiently low aircraft price, relative to cur-- 4{ =I^ rent $3 -4000 reciprocating engines. Both production rate and power level militate against the turbine. Thos, the forecast for Category I turboprop deliveries in 1988 is zero. Some subjective number could be added to this, based on the assumption that some ininital level of primary flight instruction may be performed in a turbine aircraft.

In the same table, the undisturbed 1988 forecast for Category II is 12,000 i- units. Given this demand quantity in 1988, the forecast equivalent turbine powerplant cost per kilowatt (horsepower) becomes low enough that all 12,000 units could be theoretically converted to turbine power. It was judged; how- ever, that regardless of the economic feasibility, customer acceptance 18vels would restrict the demand for a Category 11 turboprop in the 1988--93 time- frame. Thus only 80 percent of the 12,000 units forecast for 1988 were pro- jected, by marketing judgment, to be turbine powered. The Category 11 1988 forecast becomes 2400 piston powered aircraft and 9600 turboprop powered air- craft .

Because of the engine similarities in Categories III and IV, demand itera- tions for these aircraft were done with combined delivery volumes. As in Category II, the cost per kilowatt becomes low enough, given the combined f volumes of Categories III and IV, to allow total convetsion of the Category III aircraft to turbine power. As in Category II, it was felt, however, that market conditions would restrict the conversion to only 80 percent turboprop.

Thus, the forecast calls for 1100 pistidn powered aircraft and 4400 turbine powered airplanes. These were later sb^divided into Category IITU and I11P types, as previously noted, and as shown in Table T.

In Category V, where there are already turbine powered aircraft, total con-- version becomes economically and acceptance-Wise possible, and demand expands based upon the lower engine prices. Total demand in 1988 for Category V calls for no piston powered aircraft and 3000 turboprops.

The industry growth trends, with and without GATE power influences, are sum- marized in Table VIII. Overall, General Aviation is projected to grow at a simple annual rate of 4.6 percent. This rate falls into the "conservative, but probable" region of various industry projections.

Summarizing, without any introduction of a low cost turbine, the forecast is Given a lower cost GATE for a total of 21,350 piston and 990 turboprops.

turbine, this forecast becomes 7240 piston powered aircraft in 1988 and 19,570 turbines.

Market Analysis - Bell; Rotary Wing Aircraft 3.3 The market analysis for the rotary wing aircraft (helicopters) indicated that pro- gaps in terms of productivity and price exist in the currently available ducts. These gaps could be effectively filled by new designs using either GATE engines or derivatives featuring GATE technology levels. The engine the power requirements of 261 + 56 requirements summarized in. Figure 20; are .

kW engine performance levels (350 + 75 ship) fits within the GATE band, The such as fuel consumption, weights, and TDO (time between overhaul.) can . be and recently developed high technology air cooled

satisfied by both uncooled

turbine engine designs. This technology span is represented by the older T63 and the more recently developed T700 turboshaft engines (Reference, .1ANEIS ALL THE WORLD'S AIRCRAFT 1977-78). The most challenging requirement for .a new generation of small .turbine engines (GATE power class) is the initial cost of $20,000 to$33,000. One additional highly,desirable feature is

bogie

of free turbine pow-

that the engine' the torque-stall characteristics a

have

erplant.

Three different light helicopterst a single engine, twin engine and a tri--pac are envisioned, each using the same engine. The projected market is summar- a total of-4150 OEM (original ized in Table IX; 2750 helicopters requiring equipment manufacturer) engines over a 5 year time span. The potential mis- sions for this line of three helicopters include training, search and rescue, agriculture and others, as summarized in Table X.

The aircraft (helicopter) capabilities and selling prices are summarized in Table XI. The single engine helicopter has 3 seats, a useful load of 545 kg (1200 lbs), gross weight of 1271 kg (2800 lbs), range 334 kM (180 nm) at a The selling price is estimated to be cruising speed of 46.3 M/s (90 knots).

$100,000 to 125,000 (1977 dollars).

The twin engine helicopter has 5 seats, a useful load of 976 kg (2150 lbs), gross weight of 1952 kg (4300 lbs), range 834 kM (450 nm) at a cruising speed of 72 M/s (140 knots). The selling price is estimted to be $300,000 to 500,000 (1977 dollars).

The trip-pac (three engine) helicopter has 8 seats, a useful load 'of 1657 kg (3650 lbs), gross height of 3337 kg (7350 lbs), range 1019 101 (550 nm) at a cruising speed of 77.2 M/s (150 knots). The selling price is estimated to be $700,000 to 1,000,000 (1977 dollars). The range in selling price for each of the three helicopters reflects the variations in equipment and modification for the different missions summarized in Table X.

of Engine Superiority 3.4 Alternate Cycles - Identifying Domains The objective of this sub-task was to compare 1988 capabilities and require- ments for power plants of various types, and to determine the regimes in which each would play a significant role in the marketplace. In order to accomplish the task within the scope of the GATE study, published design data were researched, and various extra sources were contacted (References 1 through 4). Subsequent to completion of the task, reference 5 became availa- ble - it quantifies the other opinions on competing powerplants.

The only precise method of comparing the resulting broad range of powerplant

capabilities would be to predict the actual performance of the engines

installed in GATE-type aircraft in 1988 — this would require detailed engine and aircraft design and performance analysis. Since this depth of analysis is beyond the scope of this study, a relative value comparison, as shown in Table XII, was prepared from the literature research on engine information.

As shown in the table, the reciprocating spark ignition engine was chosen as baseline, and was assumed to be 10 percent improved in fuel. consumption, (primarily due to the anticipated emissions- driven pressure for lean burn -

combustion) . Additionally, a very conservative assumption was made that no

increment of production price will be necessary to enable the spark ignition I^ engine to meet the EPA emission targets. Reference 2 indicates that a 15 ( percent price increase could result), .L As can be seen from the table, the gas turbine is superior to all other pow- erplants in power-to - to-frontal area, as well as TBO (Time weight and power - Between Overhaul) - the diesel, is superior in fuel consumption. Comparisons V; were drawn from the assessment data n Reference 1, and combined with experi- ¢H i ence--based evaluations.

i^ The literature research, especially References 1, 2, 3, and 4, indicated that there do not appear horizon for reduction of to be any breakthroughs on the cooling or weight consultation with Beech for the diesel or rotary engines; indicated that much as 8-10 percent of current cooling losses could be as aircraft drag. This penalty against the diesel, would have to be levied ' rotary, or advanced reciprocating cell fuel consumption for engines' test comparison of aircraft performance with a GATE - derived turbine. References 2 and 4 indicate that future reciprocating engine developments will move in the 3?

direction of leaner fuel air ratios, hence increased cooling loads on the system; when for reduced this trend is combined with expected requirements weight, the ability ( TBO) greater than the to achieve time between overhaul current 1200-2000 hours will be in question. The 373-597 kW (500-800 hp) gas turbine engine has a proven ability for at least 3500 hours TBO as described in the open literature; the Task III engine studies showed that this TBO is of also achievable with Advanced Technology GATE engines. The high risk adapting a diesel for aircraft purposes would appear to indicate that company capital requirements for either development or production tooling raise seri- ous questions about the ability to penetrate the marketplace; diesel's government support could alter this assessment. In the case of the rotary engine in the USA, there is no sales or service base in existence for air- craft, hence substantial funding would be required to support the new, unfam- iliar powerplant - this could require a timeframe longer than the 10-12 years j predicted for the more familiar turbine, hence would raise serious business questions to the developing company.

^f It was concluded that the basic competition in the 1988 marketplace will be between the gas turbine and the reciprocating spark ignition engine. Primary reasons are that the application and long term for a return' on risks of ..

investment of the other two types of powerplants are expected to be too high to warrant the investment capital; this especially true would appear to be when compared to the current established status and projected improvements of the gas turbine and spark ignition reciprocating engines. Because of the private nature of such business assessments, no public projections of these considerations are known to be available.

Again, subsequent to completion withdrawal of EPA emissions of this task, regulations for the GATE fleet suggest that the costly development for lean burn combustion, or for sophisticated fuel injection and control systems may be delayed or canceled. The expected, but unquantified durability copse- quences (Refs. 2 and 4) of reduced fuel-air ratio in the reciprocating engine could lead to sustaining the current "burn for cooling" flying opera- rich tions.

9.5 Noise and Emissions Regulations Projection of noise and emissions regulations was recognized as potentially being both a major propulsion system driver, and a very speculative exercise.

United States and international environmental legal restrictions are more subject to political than engineering extrapolations over the ten year per- iod.

To bring some order to the projection of emissions criteria, references I through 4 and 6 through 10 were reviewed; the data are summarized in Table XIII which compares the EPA 1979 standards for the LTA (landing/take off) cycle for the three categories of powerplants within the GATE purview. For comparison, the QCGAT objectives are listed. In general., it can be seen the current production engines exceed the standards. Future-engines, be they reciprocating spark ignition or gas turbine, are expected to fall within at least the EPA 1979 standards.

As indicated in reference 2, extra componentry and control elements will probably have to be added to the reciprocating engine to account for emis- sions control. When combined with development requirements to move in the direction of lean burn for emissions reduction, a substantial delay could be incurred in the piston engine fleet meeting the standards. Increases of engine price are also probable. On the other hand, the rapid progress made in the reduction of gas turbine emissions, as exemplified in references 6, 9, and 10, would indicate that a smaller impact on the complexity and sales price of gas turbines could be expected.

This qualitative review can be summarized as: 1. Any of the proposed powerplants can (or could) meet the 1979 standards, given enough capital for development investment and enough time, but with possible consequent operational penalties. (e.g., In the case of the reci- procating spark ignition engine, increased maintenance due to injection and retardation controls, dr cylinder overheat.)

2.

The GATE Task II engine design studies can incorporate some of the advances made in the larger turbine engine demonstration vehicles funded by NASA and other sources - where "large" engine in this context also includes the QCGAT at 7.12 - 17.79 kN (1600 to 4000 lbs.) thrust.

3. The current reciprocating engine fleet is a mature set of engines, based on large capital investments and production tooling. The GATE gas turbine engine'would begin from a clean sheet of paper; it would of necessity include significant new manufacturing methodology to achieve the projected price tar- gets. Therefore, beginning from the research base now available in gas tur- bines, a speculative opinion is offered that if the standards were tightened for 1988, the GATE gas turbine could more easily meet them, thereby enhancing its market.

In mid-September, 1977, the aboveanalysis became moot, in view of the EPA

imposition

announcement postponing (indgfi.ni.tely) - th

y ' of the 1979 3x0 cycle

reason was the minuscule bene-

requirements it was stated that the primary

fit-Ito-cast value of the standards to the national environment.

A similar approach was taken to the comparative analysis of noise require-

ments, References 7, 8, E and 10 were reviewed and the existing General Avia- tion fleet data summarized as shown in Figure 21. The noise levels presented

are based on a 305 M (1000

are for the American General AviatioT?_Fleet,. and

as indicated in FAR

feet) fly aver measurement, corrected for rate of climb, 36. Also shown are four levels of regulation - existing FAR 361ICA©; a pro- an EPA suggestion for the

jected agreement for 1980, a draft FAA NPRM and

1980-1985 period,

The data for the current fleet indicate that differences from the 19$0 FAR

requirement this regulation could be

are small, hence it is conceivable that met should the EPA with straightfoinvard engineering development.. However,

suggestion for

1980 become law, a major crisis would exist in General Avia-

tion.

In summary, the conclusions as to noise regulations show that; 1. Potential rule making varies over a 10 dh. range in the area of interest for GATE.

a propulsion

2. This large variation could of itself drive the definition of

system. The analogy is to congressional mandates on the automobile industry

which forced catalytic converters into automobiles, even though better long

range solutions might have existed. Speatfi.cally, lower noise could, in gen-

eral, require lower tip speed propellerp, which would require higher ratia gear boxes, increased landing gear length, and outboard engine placement in s these factors has an impact, on the design, cost, stability, twins. Each of and safety of GATE-type aircraft. The smaller nacelle diameter and ease of gear ratio design favor turbine power. In the turbofan propulsion system, high levels of attenuation would be necessary in entrance and exhaust ducts,

accompanied by internal design changes for aerodynamic noise reduction; these

could have a serious impact on small engine design performance and cost.

Nevertheless, the advantage could still be with turbine power, if reciprocat- ing engine performance is seriously degraded by muffling and gearbox weight.

3. Should the lowest limits be imposed, the impact on aircraft and engine

design could substantially increase the cost of GATE aircraft, and inhibit sales. A dialogue similar to that between the automobile manufacturers and Congress/EPA could be expected.

4. In the case of the commercial airline fleet, prior NASA studies have uti- lized

a percent Return On Investment (ROI) or Direct Operating Cost (DOC)

change versus change of perceived noise; the GATE market does not respond to

these factors, as shown in Section 4.7. Most airplanes are not bought as

investments (except for the corporate aircraft), and direct operating cost is not a primary sales feature. Therefore, excessive initial purchase price increases could decrease market demand in the non-business portion of the market, e.g., Categories II, IIIU, IIIP and parts of IV.

5. During Task I1, engine detail design allows the development of more spe- cific data on the noise issue; the study plan incorporates such potential noise reducing features as the Hamilton Standard Q Fan ( tm) and 33.3 rev/s (2000 RPM) turboprop gearboxes.

3.6 Market Forecast - Task I (Summation of Findings) As previously noted in Section 3.1.1, the missions for each category were iterated in Task 11; it was discovered that the initial assumed range and velocity improvements projected for 1988 resulted in both excessive power demands and oversize aircraft in some categories. New information on pres- surized aircraft also indicated the need for iteration. Table XIV summarizes the final ( post Task I1) 1988 market projections. An assumption has been made that the 1988 date represents a mature sales market. This is recognized as being somewhat optimistic; therefore, the data more accurately represent a 1988 - 1995 time period. It can be seen that a potential, exists for upwards of , 000 annual turbine 31,000 engines per year, which represents a $220,000 engine market. This is OEM value including spares at the OEM value and based on the recommended common core design discussed in Section 5. The helicopter market is a small percentage of the total, and the assumption is that one manufacturer ( Table XIV) will capture all this market.

Under the assumption of a 50 percent market penetration by a single company, over 16,000 engines would be sold, thus justifying the assumptions made in the engine price predictions.

Two factors arise from the summary forecast. First, no penetration was pos- sible into Category I, the simplest aircraft, due to a 15 percent aircraft cost increase. On the other end of the spectrum, in Category V, the 1081 kW (1450 hp) engine size, i.e., 559 kW ( 750 hp .) flat rated to cruise altitude, is outside the scope of GATE. Therefore, it was concluded that the GATE area 265-565 hp) range of engines, and that a of interest covers the 197-422 kW ( large market potential exists.

engine features or factors which drive the Figure 22 summarizes the desirable engine design in Task I1. These features, especially the SFC (equal to piston engine), high power to weight and zero cooling drag, lead to an expec- tation of Task IT aircraft designs which are substantially improved over 1976 General Aviation aircraft -- due to GATE baseline power on its own merits.

The multi-fuel capability has both competitive and national energy conserva- tion i mplications, owing to the ability of the turbine engine to adapt to a variety of fuel types; the reciprocating engine is dependent on high octane aviation gasoline which may be in very restricted supply by 1988.

Figure 23 summarizes the accessories desired for the GATE-type engine, accom- modating both helicopter and fixed wing aircraft. Preliminary investigation of the • effect of environmental control requirements on turboprop engine per- formance indicated the desirability of a gearbox which could accept an optional or plug- - in auxiliary compressor for pressurization. Other accessory requirements were conventional Figure 24 describes the GATE interface considerations. In the fixed-wing aircraft, the most important effect on the assumed single shaft ( connected)

V

aecoFnmodate nega- turboprop was the need for a variable pitch propeller, to tive thrust on approach. In the rotary wing .case, the need for free turbine

u

torque ,characteristics was strongly stated. by Bell., and resulted in several.

Task II design evaluations.

u

Figure 25 draws together the conclusions f rom the market analysis. It is recognized that some of the specific modeling techniques used to define the engine price and the market elasticity relationship could be questioned as to absolute validity. The primary challenge for the rest of the study was to address the engine rate/price/market circular argument; to do this, turbine engine designs must be produced in Task II which reflect the engine price "bogies" shown on Figure 26.

SECTION 4.0 TASK II: TRADE-OFF STUDIES tThe Task II Study Plan for engine-aircraft trade--offs is shoran in Figure 27.

The input from Task 1, by Market segments, defines the general objectives to be accomplished.

Initially, parametric studies were run for each aircraft category, at varying temperatures and pressure ratios for all engines, and varying bypass ratios for the turbofans, all using the Teledyne CAE assessment of the 1985 compo- nent state--of-the-art. The results of these parametric studies defined areas of concentration to be addressed in the conceptual engine layouts.

L11 These layouts incorporated inputs from the TCAE component and configuration data bank, large engine and other recommended technologies, and sub-contrac- tor inputs such as the Hamilton Standard Q-fan(tm), advanced propellers, and fuel controls. These engine layouts were then iterated with cost and perfor- mance analyses, using in-house manufacturing and vendor data bank projections as criteria In parallel with the engine studies, Beech Aircraft provided conceptual air- craft layouts and their evaluation, as well as parametric performance ana- , lyses around the point design, using baseline engines. Trade-offs were accomplished with the competitive engine layouts using the aircraft sensitiv- ity calculations to assess cost.:, performance, and take--off gross weight devi- ations from the baseline.

^; From the aircraft-engine synthesis information, Life Cycle Cost Analysis was accomplished, and an optimum engine selected for each category of aircraft.

The task output covers power and SFC -vs-- cost trades, a description of the optimum engines and their requisite features to meet the challenges laid down r, by Task I, the aircraft layouts and their performance/cost assessment, an analysis of the benefits of the improved aircraft compared to the current General Aviation fleet, an assessment of the applicability of various tech- nologies and their worth to the GATE concept, and finally the systems cost analysis and environmental impact.

4.1 Parametric Performance Analysis A parametric cycle analysis of turboprops, turbofans and turboshaft is the starting point to determine the 'optimum" engine configuration(s) for the GATE missions. This analysis included the effect of reducing component effi- ciencies in the small (GATE) flow sizes, turbine cooling?, and cycle pressur ratio.

4.1.1 Baseline Analysis Parametric performance was calculated for turboprop, turboshaft, and turbofan configurations, using projected levels of efficiency capability for the 1988 time frame.

accomplisl. for three mission conditions: helicopter - sea The analysis.was - 3048 m' (110 Category 12 and IIIU (Unpressurized) level 56.6 m/sec KTAS): II:IF and TV - 5486 m .(10,000 ft) and 92.6 m/sec (180 knots); and Category In the latter two cases, an average 123.5 m/sec (240 KTAS).

(1. 8,000 ft) and the velocity and mission was used, because there was insufficient spread of altitude conditions to make separate analyses worthwhile. Table XV'provides the matrix of efficiency assumptions used in this parametric analysis, and in the determination of sensitivity values of, for example, power and SFC to component efficiency levels. The sensitivity values are desirable to allow later iterations of performance with the effect of selected component configurations. These initial calculations are purely parametric, to outline trends.

Conventionally, this type of analysis is done by assuming constant efficiency levels, then varying the major cycle parameters such as pressure ratio and temperature at the critical mission point (assumed to be cruise, in this study). Optima are then deduced from the shape of the curves; sensitivity values are calculated for small changes to the input efficiency assumptions, and priority design drivers are determined from the results.]

4 -+ component During the time frame of the parametric analysis, the parallel evaluations (described in Section 4.3) were producing results which indicated' that the efficiency levels of components in the small airflow category of the GATE problem be constant across the range of pressure statement would not ratios indicated Therefore, the analysis was repeated at key in Table XV.

conditions with the analytically determined variable efficiencies of Sections data output of 4.3.1 and 4.3.3 using sensitivity values determined from the IIIP and IV Table XV. The results are shown in Figure 28 for the Category mission statement.

mission statement, and in Figure 29 for the helicopter pressure :-^ They show a trend of decreasing specific power with increasing pressure ratio. SFC is ratio, and an optimum, or minimum SFC at a specific seen to be minus 1 percent over a substantial relatively constant (plus or to the ability to meet low pressure ratio range). This latter trend leads number of component arrangements, as discussed in SFC requirements with a Section 4.4. In each case, it is evident that the analytical or "real'' effi- ciencies tend to reduce the level of optimum pressure ratio from the 15 to 20 I range to the 9 to 12 range.

j 4.1.2 Sensitivity Studies paid in the determine where priority attention must be In order to help F..

engine component choice and design assessment, sensitivity analyses were run at two different flight conditions, one for Category IV and one for the heli- show the copter, as indicated in Figure 30. The sensitivity coefficients per percent change of compressor percent change of specific fuel consumption and turbine The curves show that the efficiencies, versus pressure ratio.

turbine overall a higher impact on engine performance than efficiency has does compressor However, both are important, in that they are efficiency.

consumption per percent` above 1.0, i.e., more than 1 percent change of fuel The figure also shows that both the change of either component efficiency.

higher pressure ratio lower altitude mission (helicopter engine) and the These sensitivity coef- cycles are more sensitive to changes in efficiency.

of component arrangements ficientsaid in determining priorities for choice 1s ^J

I

in detail engine design IT T 4.1.3 Coolie Effects cts It is commonly accepted that increasing turbine inlet temperature is benefi- cial to most gas turbine cycles; in the case of the turboprop cycle it is especially beneficial. Therefore, considerable attention was paid to methods of achieving higher turbine inlet temperatures than are currently opera- tional. Because of Teledyne CAE's considerable experience in the cooling of very small turbines with blade spans between 6.35 and 12 . 7 mm (0 . 25 and 0.5 inches / ( Reference 11) and the low efficiency test results of that experi- ence, the input assumptions used for the high temperature analysis were care- fully evaluated.

These assumptions are summarized in Figure 31 as an effect on turbine effici- ency and the cooling bypass bleed required by increasing turbine inlet temp- erature. Figure 31 shows that if an all-axial turbine had been assumed for the design, the low aspect ratio and thick blade shape problems ( resulting from the difficulty of cooling small blades) would reduce turbine efficiency L-t capability by 6 to 8 percent below design requirements and the parametric ,-^ data assumptions. This same figure also shows that if the radial portion of 1 u a radial-axial configuration were to be cooled, an efficiency loss of between 1 percent at maximum temperature ratings of 1421 degrees K (2100 degrees F) and 5 percent at 1643 degrees K (2500 degrees F) would be incurred. The com- ponent analyses showed that the axial element of a radial-axial turbine com- plement would not have to be cooled, because of the large temperature drop through the radial turbine ( see Section 4.3.3). The study baseline rotor is radial, and is uncooled to a takeoff turbine inlet temperature level of 1504 degrees K ( 2250 degrees F).

^n The right hand side of Figure 31 indicates the equivalent radial turbine bypass bleed as a function of temperature and cycle pressure ratio. The baseline value for a high tip speed (uncooled rotor) radial inflow turbine, is 2 percent of compressor exit bleed required to cool the turbine inlet nozzle and the shrouds. Most of this flow is recovered for use, in the rotor, hence does not represent a large cycle loss. It is known from design experience that heat flux and Nusselt number both increase as the turbine pressure level is increased, thus, higher levels of cooling bleed are required for the 20:1 cycle than for the 9:1 pressure ratio cycle. It is evident that an important input to a GATE engine design is the ability to attain an uncooled radial inflow turbine rotor at temperature levels as high as possible, providing zero reduction of turbine aerodynamic efficiency, and a requirement for approximately 2 percent bypass bleed ( at the 9 : 1 pressure ratio).

Figure 32 shows the resulting effect of increased turbine inlet temperature on specific horsepower and specific fuel consumption, as temperatue is raised at the Category IIIP / IV flight condition of 123 M / s (240 KTA.S) at 5486 m (I8000 ft). Preliminary engine and aircraft matching analysis had indicated that an approximate difference of 167 degrees K (300 degrees F) exists bet- ween the cruise inlet temperature and the maximum thermodynamic power turbine inlet temperature for this mission ( see also Section 4.2).

Performance is shown for the upper and lower limits of cycle pressure ratio 9:1 and 20:1. At the 9:1 cycle pressure ratio, two levels of bleed ("optimistic" and "conservative") loss a ..ssumptions are shown., to illustrate the range of consequences of cooling the small 1.19 kg/s (2.6 lb/sec) flow rotor, if required. Table XV1 illustrates the losses used for the analysis•, the values range above and below those shown on Figure 31; the 1°conservative" values represent upper limits, based on previous 'Teledyne CAE design experi- ence and materials. The power (Figure 32) is seen to increase substantially with cruise inlet temperature. Coaling effects are shown to represent a minimum of 3 and a maximum of 10 percent increase in equivalent specific fuel consumption, hence an objective of the GATE engine design is to minimize these losses.

4.1.4 Turbofan Performance Evaluation

The matrix: of data assumptions used to do the turbofan parametric analysis is

shown in Table XVII. Efficiency assumptions are consistent with advanced technology GATE levels used throughout the study. The calculation was made at Category I q flight conditions because these offer the- best specific fuel consumption opportunity within the Category 11 - IV range for a turbofan application. If the analysis had indicated a turbofan to be superior at these conditions, additional calculations would have been made at the lower flight speeds and altitudes. The performance resulting from these parametric Sa assumptions is summarized in Figure 33 as specific fuel consumption -vs- bypass ratio for a range of core pressure ratio and temperature assumptions.

In each case, fan pressure ratio was optimized for the cycle, thus the solid and dashed curves 'represent a range of turbofan SFC capability for the core technology levels assumed in Table XVII. It is seen that the thrust specific fuel consumption of the turbofan is. 20 to 30 percent higher than the previ- ously calculated range of turboprop specx£ic fuel consumption.

In order to offer the most optimistic comparison of the turbofan and the tur- boprop, the design duct Mach number was reviewed in an attempt to reduce the fan duct loss for the loss-sensitive high, bypass ratio cycle. The casing of the assumed typical Category IV turbofan was increased by 43 . 2 mm (1.7 inches) ( approximately 10 percent of baseline engine diameter) and the losses recalculated at a bypass ratio of 8. The-duct loss was reduced from 4 to 2 percent; the result was still an 18, percent increase of specific fuel con- sumption over the turboprop.

It was therefore concluded that on a specific fuel consumption basis alone, the turbofan was not competitive; the configuration is not excluded from further consideration, depending on the final study results and the applica- bility of GATE power to the market categories. A turbofan could well be con- sidered as a fall- - out derivative of a successful GATE turboprop engine, for special markets requiring either higher speed than was assumed in Category IV, or ,a more attractive aircraft configuration.

4.1.5 Parametric Performance Analysis Conclusions The parametric cycle analysis Zed to the following observations: 1. A turbofan is not competitive (on an SFC basis only) in Categories II -

market

1V. A derivative of a successful turboprop could provide a special for a high performance airplane.

2. Turbine cooling is high risk in this size of powerplant, and offers a

very limited payoff. Teledyne CAE foresees no breakthroughs in the industry

ability to achieve a small cooled axial flow turbine under 1.3 kg/s (3

lb/sec.) with a suitable efficiency. Manufacturing techniques, thin-wall all militate

materials property degradations, and heat transfer limitations

against the expected production price of an effective design. Additionally,

a small cooled turbine would run counter to a major requirement of this study

- prices. This would be especially

the need to approach reciprocating engine true in comparison to the uncooled radial turbine at 1504 degrees K (2250 degrees F) take -off temperature.

3, high pressure ratio cycles are more sensitive to flowpath efficiency and are prone to flange leakage.

efficiency.

4. High temperature cycles are less sensitive to flowpath

Design-derived (real) efficiencies estimated for GATE-type components 5.

decrease optimum pressure ratio for minimum specific fuel consumption.

Therefore, the optimum GATE engine should have a medium pressure ratio (range of 9-12) and the maximum allowable uncooled take-off temperature: 1420-1504 degrees K (2100-2250 degrees F).

4.2 Same Engine Performance Studies A computerized turboprop engine performance model was constructed to allow additional evaluation of component design priorities (matching, surge margin, definition of ratings) and better integration with ongoing Beech aircraft design studies.

The model was constructed for a 9:1 pressure ratio cycle (C9 configuration) connected shaft turboprop (Fief. Section 3.1.2). Maps for the compressor and turbine were synthesized from available Teledyne CAE and open literature data.

Prior to calculating wide-range performance, investigations were conducted to evaluate the effect of turbine back pressure on the engine design and perfor- mance. It was hypothesized that an increase of back pressure at the turbine exit would have a small effect on ESFC, but could reduce turbine stresses significantly: at a constant flow and radial turbine exit Mach number, the annulus area will reduce with increasing back pressure. This in turn reduces exducer AN2 (annulus area x square of RPM), which is a direct measure of exducer stress.

Figures 34 and 35 summarize the results of the analysis for Category IV flight conditions and sea level static (SLS) respectively. They verify that the turbine back pressure can be increased beyond conventional practice; as an example, increasing it to 1.25 times ambient, increases specific fuel con- sumption by only 2 percent. This conclusion holds with the three engine speeds shown on the figure, and over most of the applicable operational power

spectrum. A tradeoff therefore exists between specific fuel..consump tion and

either

turbine durability (lower specif3a fuel consumption . , higher stress) or

between specific fuel consumption and a higher cost turbine configuration

(radial plus axial, to maintain suitable stress levels). Subsequent prelimi-

nary engine designs assumed the 1.2$ x ambient value, to maintain . the lowest

cost turbine configuration.

The figures also indicate that the falloff of efficiencies resulting from the

assumed component characteristics actually results in increased cruise spe-

cif:ic fuel consumption

above design corrected speed.

Figure 34 shows performance for three engine speeds: 100 percent, 95 percent, and 90 percent of

design mechanical RPM rating. At the Category TIIP or IV

flight condition, compressor face temperature = 262 degrees K (471.2 degrees R), they correspond to corrected speeds of 104.9, 99. 7, and 94.7 percent of

design respectively. At a typical cruise condition of 1.068 kN (240 lb)

thrust, Table XVIII shows the performance at each of these conditions. The

best cruise performance, as to required temperature (durability) and specific fuel consumption is obtained at approximatly 100 percent , of design corrected speed. Tc maintain best climb and cruise SFC, it was concluded that this

particular configuration should be controlled to operate at constant mechani-

cal, speed at inlet temperatures over 288,.5 degrees K (60 degrees F), and con- stant corrected speed at inlet -temperatures below this value. A 2.3 percent maximum thrust reduction is implied by this rating method -- from 1.530 to 1.495 kN (344 to 336 lb) at 1504 degrees K (2710 degrees R) on Figure 34, but the 6 percent specific fuel consumption improvement and 38.8 degrees K (70 degrees R) reduced temperature make the trade-off worthwhile. Additionally, the engine would run at less than 100 percent mechanical design speed at all inlet conditions less than standard jday, with a consequent durability improvement relative to a constant RPM engine.

Detail performance calculations were then completed at a jet nozzle area of .01129 m (17.5 square inches), representing a judgement compromise. Subse- quent studies should further evaluate this rating method as a means to opti- mum component design. The rating method also affects engine installation features: 1. Excessive jet velocity due to a high pressure ratio could cause taxi area

erosion in a

single engine aircraft if the exhaust is under the aircraft nose.

2. The effect of rating method on the match of the connected shaft engine in taxi or acceleration mode might be significant. It is desired to warm up and taxi at minimum propeller speed for noise and erosion attenuation, yet to have the instantaneous prop response of the connected shaft engine. This requires the ability to decrease engine RPM to 60-70 percent of design, with- out endountering surge - a difficult task for a connected shaft engine.

The engine performance was calculated for a nominal 1.19 kg/s (2.62 lb/sec) corrected design flow (approximate match of anticipated Category IV aircraft peint design horsepower) over the range of flight conditions. Typical equi- valent specific fuel consumption and thrust data

-vs- power level are shoe

in Figure 36 for the Category TV cruise and maximum altitude conditions

It

t

a I F

(solid and dashed lines respectively). For subsequent scaling of the i configuration to a specific aircraft design (Section 4.5), a maximum cruise rating scaling point (circle) was chosen at 1328 degrees K (2390 degrees R) as.a balance between weight, durability and SFC.

W.

L The analysis also showed that an initial assumption - that cabin pressuriza-

tion air if bled from

compressor discharge pressure (CDF) - resulted in 6-9 percent power and SFC loss, due to the small airflow of the engine. It was therefore decided that all accessory power would be taken from the gearbox, including a mechanically-driven.compressor. This reduced the pressurization SFC loss to between 2 to 3 percent.

t''.

`=I' The studies thus provided background data for the operation of a typical g p YF advanced technology GATE engine as an aid to outline both a need and a direc- tion for future, in-depth, evaluations.

4.3 Component Analysis This section describes the development of candidate component data for input to the trade-off studies to define an optimum engine type for each of the market segments identified in Task I. The rationale for development of the component data is as follows: 1.

The maximum performance potential of compressor types know to be applica- ble to the GATE market was determined. By experience, it is known that cen- trifugal and axial-centrifugal compressors suit the 0.454-1.361 kg/s (1-3 lb/sec) market place, and that all-axial compressors do not. For this rea- son, single centrifugal, twin centrifugal, and axial-centrifugal combinations (from one to three axials) were chosen, and their limits and domains of supe- riority identified.

2.

The impact of these compressor configurations on engine aero/mechanical design and structural integrity was assessed to determine best choices for detail engine incorporation.

3.

Four configurations were chosen as representing the required trade-off potential. One of each of the configurations was selected: single centrifu- gal, one axial plus 1 centrifugal, 3 axial.s plus 1 centrifugal, and 2 centri- fugals, covering the pressure ratio range from 9 through 20 (as determined from the parametric cycle analysis above).

4.

The turbine requirements, their temperature capability and their perfor- mance potential, were evaluated in reference to the speed and pressure ratio determined from the compressor study. Both radial and radial-axial combina- tions, for connected shaft and free turbine configurations, were evaluated.

5. A short combustor evaluation was conducted to determine whether the com- bustor was a critical component, or could be drawn from ongoing technology programs.

4.3.1 Compressor Component Studies A series of nine typical compressors, covering the range of pressure ratio i from 9 through 20, all applicable to small engines, was, analyzed as summar- ized in Table XIX.

Each of the configurations was evaluated using standard designn -techniques, plus efficiency algorithms. For -the axial compressor_, the model included single and multi-stage aerodynamic and tip speed loading effects., as well as size effects; centrifugal con'P'1.gurati,on analysis included specific speed, backward curvature, pre -whirl factors and size effects.

The algorithms were then used to extend the parametric analysis to define envelopes of maximum efficiency -vs- pressure ratio for each type of configu- ration, thereby determining domains of component and configuration superior- ity. In each case, usable (operating line) efficiency was assessed.

Figure 37 presents the results, in terms of compressor polytropic efficiency -vs- pressure ratio, for various configurations of compressor, to allow^com- parison with o Existing state-of-the-art of small compressors., compiled from Tele- dyne CAE and open literature data.

o The large engine state-of--the- a-tt, compiled from open literature data and recent NASA-sponsored compressor studies on the 198X energy effi- cient, all-axial transport engine..

This figure verifies the general curvature and level of the assessment der- ived from the algorithms.

From this matrix of compressor designs., four were chosen for preliminary engine designs as representative of applicable trade-off data. The four com- pressors are described in Figures 38 through 41. Figure 38 shows a 9:1 pres- sure ratio single stage centrifugal, the Jsdmplest configuration. For maximum flexibility of match (See Section 5.0), variable inlet guide vanes are included. The radial diffuser is configured for maximum diffusion efficiency of the supersonic inlet Mach number. 'A separable inducer is shown as one method of efficiently handling the transonic tip Mach number resulting from the design.

Figure 39 shows an axial--centrifugal, des-i.gn at 11.3:1 pressure ratio. The

stage pressure ratio split was chosen from the design algorithms for each element to maximize overall efficiency. The small size (flow) effect has been taken into account for the axial stage, at 82.2 percent adiabatic effi- ciency. A transition duct is i.ncorporatied, both to eliminate inducer vibra-

tory effects from the stator wake, and to allow a sufficiently low inducer

hub-tip radius ratio for maximum centrifugal stage performance.

Figure '40 shows a three--axial, one centrifugal design, at 15.0-1 pressure

ratio. The design philosophy is an extension of the approach for the 11.3:1 pressure ratio configuration. Of partictil.ar significance is the blade height in the third stage axial, which is only 10.2 mm (0.4 inches) at the 1.19 kg/s (2.62 lb/s) floc; size - at the lower level of manufacturing feasibility.

Figure 41 shows the two stage centrifugal design, at 19.9:1 pressure ratio..

Each 'stage is optimized for efficiency. Significant results evident from the design are the diameter - greatest of all the configurations, the "swan's neck" transition duct, and the very narrow (2.51 mm/0.099 inch) second stage diffuser width.

The compressor study showed that the GATE compressor design challenge is approximately plus 2 points in compressor efficiency. The program plan des- cribed in Section 6 addresses this need via: 1.

In general, better clearance control., using improved deflected housing design techniques, and improved abradable coatings.

2. Improved design system - a better understanding of secondary and leakage flows in both types of compressors is required.

3.

In the axial case, more effort is required to define the efficiency capa- bilities of low aspect ratio, highly loaded, configurations.

4.

In the centrifugal, an intensive effort is required to develop better knowledge of the transonic inducer.

5. Again in the centrifugal unit, backward curvature must be extended to the 9-10 pressure ratio regime (this will not only increase efficiency level, due to reduced Mach number, but also tend to move the efficiency islands away from surge, into the usable match regime).

6. The advanced GATE engine design does not restrict radial diffuser depth, hence limits of diffusion may be relaxed.

To meet GATE performance and stability needs, the compressor effort outlines a research challenge and a path to meeting that challenge.

4.3.2 Combustor A reverse flow, vaporizer-plate combustor was chosen as baseline, based on the promise showy ► by the configuration in current TCAE programs for reduced cost and reduced emissions potential, at excellent performance. A typical configuration is illustrated in Figure 42. It is a conventional reverse flow design, with a unique fuel injection scheme. A cylindrical annulus is used as an aid to spreading fuel from discrete injection points. Heat transfer across the vaporizing plate is used to pre-vaporize the fuel, which is then injected into the primary combustion zone via the turning action of the end cup ., A performance analysis was performed for the 11.3, 15, and 20:1 pressure ratio engine designs at a cruise condition. The results were based on prel- iminary engine layout combustion volumes, and indicated that all intensity, dwell time and loadings Factors were within current design practice. This is the expected outcome of the choice of compressor pressure ratio and diffuser depth,.

This conservatism is evidenced by the summary performance shown in Figure 43.

A conventional plot of efficiency -vs- aerodynamic loading is shown, with i1 ; Li successful, rig and engine data points superimposed. It is seen that both starting and steady state aerodynamic loadings are lower than previously demonstrated rig and engine data. k' It is therefore concluded that the combustor is not a critical research chal- lenge or limit to GATE progress. A well-structured but straightforward development program is expected to provide the required performance, durabil- ity, and gradients.

{ } 4.3.3 Turbine Component Design Prior studies have indicated that engine cost, and as an input to it, maximum allowable uncooled temperature, are major drivers in the applicability of GATE turboprop engines.

For these reasons, it was assumed that the simplest 4f uncooled turbine configuration would best suit GATE engine design features.

Teledyne CAE has prior experience ti in high speed, high pressure ratio Y p s P g p P g p radial inflow turbines on small_ engines. This experience was at a 609.6 M/s (2000 feet per second) tip speed and 1227 degrees K (1750 degrees F) maximum turbine inlet temperature, on a 89.5 kW (120 hp) turbogenerator set develop- ment for Ft. Belvoir-Mobility Equipment Research & Development Center. This ) experience, combined with the recent work of two recognized authorities in the field of radial. turbomachinery, O.E. Bal,je and H.J.

Wood (references 12 and 13) indicated that considerable advancement in this technology could be predicated, given a suitable research and exploratory development program and the availability of modern nigh strength material . The high tip speed of these designs results in turbine rotor temp- tip and wheel relative metal eratures within acceptable levels for uncooled operation.

This capability, combined with

the results of the parametric performance ana- lysis shown in Section 4.1, showed the single stage radial turbine to offer the simplest configuration, with a substantial_ SFC benefit over an air cooled axial turbine arrangement. Tradeoff studies shown below and in Section 5.0 verify the initial analysis.

A radial inflow turbine, running at tip speeds greater than 701 m/s (2300 ft/sec), and thereby capable of running uncooled at temperature levels up at 1532 degrees K (2300 degrees F) is a prime candidate for baseline engine design studies, Figure 44. The turbine aerodynamic and structural evaluation studies therefore proceed from this assumption as follows.

The turbine design requirements are established by the compressor component study results. These !

results establish the turbine work load and speed.

Available computer programs were used to define the preliminary design of j turbine configurations suitable for driving the four candidate compressor designs described in the previous section.

Figures 45 through 48 describe the connected shaft turbine configurations.

The analysis showed that the pressure ratio centrifugal configuration, 9:1 C-9 could be driven by a single radial inflow turbine running at 759 m/s (2490 ft /cec, tip speed. At this tip speed, the blade metal relative temperature at the rotor tip was approximately 1258 degrees K (1806 degrees F'), hence the rotor could be run degrees K (2250.degrees F) uncooled, at . 1504 assuming advanced uired either one or two materials. The other designs req ax i al stages in addition to the highly loaded radial, stage.

Preliminary velocity triangle analysis showed rotor design exit relative velocity to be low near transonic M- 0.9) diameter. This level leads to the ( at. the mean assumption that some aerodynamic rig work is required in the follow-oil program.

Initially, an arbitrary work distribution was assigned to the radial and axial components; it resulted in an assessment of efficiency potential for all configurations, at the 1.19 kg/s (2.6 lb/sec) size, of 88 ^ 0.5 percent.

t However, this arbitrary work distribution resulted in a significant difference in temperature capability of the various designs; it was determined that the lower tip speed of the radial turbines for the AM 1.3, AAAC15, and CC20 configurations would result in a reduction of approximately degrees K (120-140 F), in maximum rated (thermodynamic) 87-78 degrees performance of these engines, temperature capability. This degrades the but only slightly in specific primarily in specific power (hence in size), fuel consumption.

In conducted of the turbine aerodynamic Task III, a re-evaluation was testing. The work was redistributed to increase the loading on the radial component, and decrease loading on the axial component. It was found that the efficiency was reduced by only 0.2 points, but the temperature capability (2250 degrees F) baseline.

increased back to the original 1504 degrees K D etail calculations showed that as the work load and tip speed were raised on the relative temperature at a given radial element (to reduce the blade turbine gas temperature), specific speed, which are basic measures of the became slightly efficiency, remained almost constant. The axial turbine work capacity) than the smaller (i.e., lower wheel speed, hence lower configurations shown on the preceeding figures, but without efficiency change.

In hindsight, it was concluded that the AC11.3 engine could have been rated at a substantially higher specific power and a slightly reduced specific fuel consumption, thereby providing a smaller, lighter engine at a reduced cost.

These results were incorporated into the Task III common core evaluation, but are not included in the engine design and life cycle cost analysis described below.

^. 27 i Parallel studies were performed for a free turbine version of the AC11.3 engine, as summarized in Figure 49. In this case, it was concluded that the reduced work load on the radial turbine component (which now drives only the gas generator compressor), did result in a reduction in temperature capabil- ity of over-92 degrees K (165 degrees F). This would provide a free turbine variant of the turboprop concept, but at a lower power capacity and slightly higher SFC level. The temperature-tip speed trade-off study was also con- ducted for this configuration during Task 111. It indicated that at they fixed level of turbine work, a reduced temperature capacity would result as the tip speed was arbitrarily increased in an attempt to reduce blade rela-

tive temperature. Efficiency would also fall off due to the bad mismatch of j!

increased turbine Parsons number (U/Co = tip speed/isentropic "spouting" vel- ocity) and increased specific diameter, Figure 50. The isentropic spouting velocity is defined as: Co 2gJc His Cohere: g = Gravitational Constant Jc = Joules Constant His = Isentropic Enthalphy Drop 4.4 En ine Configuration Layouts The component characteristics from Section 4.3 were combined into out- line layouts and performance evaluations of each of the types of the engines - at a constant airflow. The resulting power output therefore represented the effect of component efficiency, pressure ratio choice, and temperature capability on the configuration and its ability to produce high specific power or low fuel consumption (SFC).

To establish a common baseline, each of the outline layouts was adjusted to a constant Category IV power for subsequent scaling to each market cate- gory. In parallel with the outline layouts, scaling limits and size effects were determined.for each of the comp anent configurations. These effects were used to modify power, SFC, weight, and cost for the lower power levels of Categories II and III.

4.4.1 Performance Analysis The component assessments of Section 4.3 were integrated into a series of computer calculations to determine the cruise and sea level static (SLS) max- imum thermodynamic performance of the four engines at a constant 1.19 kg/s (2.62 lb/sec) flow, which was baseline for scaling studies to match antici- pated aircraft designs. The element input data is summarized in Table XX.

The data represents the results of the component studies of the previous sec-} tions ;.an increase of I percent in propeller efficiency was assumed, as a result of discussions with Hamilton Standard on future propeller development trends,_ Tle cruise performance is summarized in Table XXI for the four engine configurations being used to . illustrate engine design.tradeoff poten- tial relative to aircraft performance. The AC11.3 configuration is'seen to offer 2-7 percent lower SFC than the-other models, and the C9 up to 35 per- cent higher specific power than the other models. These differences are sig- nificant in relation to anticipated.effects on the aircraft. designs, The performance points were thenrun.at SLS conditions at the turbine inlet temp- eratures corresponding to maximum rating determined by the initial tradeoff studies. The results are presented in Table XXII. These estimates were com- pared to lapse rates for existing, larger engines, and found to be represen- tative.

A baseline set of performance data was thus established for each of the four configurations, at a constant design airflow. This allowed scaling of each engine to a constant power - for comparative sizing, and to a match power for each aircraft design (Section.4.6).

The components for each engine were then re-analyzed to establish performance degradation guidelines in scaling from the baseline size to power levels as low as one-half of design.

In the compressor cases, the design algorithm on size effects was utilized, with results as shown on Figure 51 (left panel.). Two conditions were evalu- ated: constant absolute clearance (solid line) and constant percent clear- ance (clearance/height constant, dash Line). The results, as expected, showed a sensitivity of each design to increased clearance and reduced size.

Axial elements increase this sensitivity, to the point where manufacturing limits prohibit a scale of the AAAC15 configuration below 1.04 Kg/S or .00116 m (2.3 lb/sec., or 0.4 inch) blade height by judgment.

The right hand panel shows similar effects for the radial turbine, but at slightly lower degradation levels.

The engine performance assessments were modified, using these effects, as the aircraft engine match power levels were determined.

4.4.2 Mechanical Design and Cost Analysis Schematic layouts were prepared for the four candidate engines discussed in the performance analysis section, as illustrated in Figure 52. They are sized for the same power output - 365.5 kW (490 hp) at sea level static rat- ing. Design numbers 2010, 3010, 4010, and 5010 are assigned to each configu- ration as an aid in tabulating subsequent data. These numbers are represen- tative of an engine configuration and are retained when the engines are sca- led to different sizes.

A complete engine cross section for the 2010 design (Figure 53) was prepared and a detail weight analysis made. The other engine weights were calculated using' this for a base, and modified to reflect the effects of the different flow path geometries.

The turboprop installation drawing is shown in Figure 54 and the basic scal- ing data shown in Table XXIII. Two sets of data are presented, one at equal power and the other at equal airflow. The flowpaths were originally sized for equal air flows, and subsequently scifl.ed to a constant power. The reduc- tion gear sizing was maintained constant for the 365.5 kW (490 hp) size in both cases.

The simple cantilevered rotor suspension of design 2010 is dependent upon be- 4- ing able to achieve adequate critical speed margins. A critical speed analy- r ^I sis was performed and the results summarized in Table XXIV. The estimated shaft support stiffnesses, both front and rear, of 175 mm/m (106 lb/in) uses the combined bearing and housing spring rates, and results in adequate criti- cal speed margins. The second critical speed, 1530 rev/s (91084 RPM) pro- vides 30 percent margin over the maximum rotor speed of 1148 rev/s (68,900 RPM) and the first critical speed, 217 rev/s (13024 RPM) occurs more than 60 sf percent below the engine idle speed of 804 rev/s (48230 RPM).

The ability of the radial turbine rotor to operate uncooled at high turbine inlet temperatures is an important factor in achieving good performance and low cost. Arnold and Balje (Reference 12) discuss the temperature and expan- sion ratio potential for uncooled radial turbines in the range of the GATE design 2010. A preliminary analysis of the turbine design was performed to verify that the rotor is a worthy candidate for engine development. The results of the analysis are summarized in Figure 55. Two different blade area taper ratios (ATR) were drawn to confirm that the geometries were attainable. The stress rupture life at the cruise rating was bracketed bet- ween 3009 and 10,000 plus hours for the ATR range of 16 to 31. The stress rupture life at the maximum temperature and maximum minus 28 degrees K (50 degrees F) provides a reasonable starting point for a detailed design. Two aspects that can further improve the rotor life are the maximum temperature rating and improved material properties. The aircraft operating dharacteri- sitcs result in essentially a flat rated engine and will not require the indicated high temperatures at take-off. Therefore, the desired maximum .r temperature will be established by a function of maximum altitude capability versus the cruise altitude and the desired maximum to cruise power ratio. A detailed analysis of this was beyond the scope of this study. The life ana- lysis was based on current, well-characterized equiaxed ICI-100 to give a high confidence level to the results; improvements such as directional solidifica- tion, advanced materials, and fabrication methods will further enhance the material capability, and increase the integrity of the design. Several advanced materials were screened for the application, and reserved for detail evaluation of cost-yield-strength tradeoff in Task I of the follow-on program (Section 5.0).

The relative cost summary for the four basic engine designs is summarized in Table XXV, and OEM (Original Equipment Manufacturer) cost in Table XxVI.

Figure'56 summarizes the effect of increasing cycle pressure on engine cost and engine power to weight ratio. The increased cycle pressure ratio pro- vides a modest reduction in fuel consumption when the component efficiencies are adjusted to reflect the small sizes, as discussed in Section 4.4.1. The higher cycle pressure ratios require more components, with a resulting increase in relative cost. The four basic engine designs, shown in Figure 56, require an increase in the number of turbine stages, from one to three, CPR The bottom as the (Cycle Pressure Ratio) increases from eight to twenty.

Line of the vertical band represents the minimum number of compressor stages to achieve the CPR - one centrifugal for 9:1, and two centrifugal.s for 20:1.

The to line represents a parallel band drawn through the 4010 design which has three axials plus one -centrifv al compressor. P g P The decreasing , power to weight with increasing.CPR is shown for three engine power sizes ranging from 365.5 ` k1f (49 . 0 hp)' kW (250 hp). This band is based on scaling the engine configurations using the basic size data and scaling exponents from Table:=IT.

4.493 Alternate Configurations special Two different p ropulsion configurations were evaluated to meet the s g.

P P P GATE problems imposed by potential low noise requirements and helicopter applications.

T!'" Low'-Noise Requirements: The Hamilton Standard Division of United Technolo- gies provided preliminary design information on the potential integration of the Q-FAN(tm) with GATE technology. If the installed fuel consumption of the high bypass, variable pitch fan were to prove competitive, the low noise sig- nal of the unit would provide a desirable powerplant. The design is summar- ized in Table XXVII... The configuration is compact, but the cowl drag results in an unacceptable installed fuel consumption at flight speeds of over 77.2 m/s (150 knots). This information was verified by Beech, who had conducted a more thorough study in 1975, and arrived at the same conclusion. This is L'L further substantiated by the Metzger and Worobel (Reference 14) study of Q-FAN(tm) propulsion systems; they show that a 289 kW (387 hp) core engine is required to meet the cruise thrust of a conventional propeller engine with a 213 kW (285 hp) reciprocating engine.

It was, therefore, concluded that unless noise becomes an overriding consid- eration, the Q-FAN (tm) is not a GATE candidate, Even then, a more detailed " J comparison with the p quieter GATE fleet using a 33 rev/s (2000 RPM} propeller would be necessary.

r^ Differential Turbine: Bell Helicopter Textron reflected the strong feelings of the United States helicopter industry against an form of connected shaft P Y g y engine (with a clutch to the main rotor): the torque characteristics are unsuitable, a fixed shaft installation requires a higher power than a free turbine, and even than surge when high cyclic pitch is the engine can demanded.

Different engine configurations were reviewed, using the basic fixed wing GATE simplicity and low cost to respond to this need - even though the heli-

copter represents only 5% of the total. market. A different; ^l. turbine engine

was chosen as a potential candidate. The power characteristics of this type engine 'are shown in Figure 57. It operates as a connected shaft engine to the right of the lockup line; to the left of this line the ORC (overrunning clutch) in the output drive train allows the turbine to slow down and run at the compressor. Differential gearing divides the a differential speed to torque between compressor, turbine and output as prescribed by the design ratios. During operation with the engine at part power and initially in the .Lockup region, when the cyclic pitch is increased, the OP.0 releases and the engine moves nearly vertically along the output speed Line, and crosses the "lockup" line to provide increased power in a regime where a connected shaft engine would surge.

Schematic layout designs for a differential turboprop (Figure 58) and a dif- ferential turboshaft (Figure 59) illustrate the basic concept, size, and weight for a power output of 365.5 kW (490 hp). They use the design 2010 (C9) compressor, combustor, any 1 turbine. The turbine drives through the com- pressor bore to a differential gearset that drives the power output shaft and the centrifugal compressor, The relative cost (for equal weight) of the differential turboprop and turbo-- shaft compared to the baseline, is summarized in Table XXVIII. The signifi- cant parameters; cost, weight, and SFC are compared in Table XXVIX. The dif- ferential turbine designs are significantly heavier and more expensive than the single shaft turboprop. The increased fuel consumption is due to the increased gear loss to drive the compressor. The cost and weight penalties of the differential design lead to the conclusion that it is not a GATE fixed wing aircraft turboprop candidate.

The cost and performance applicability of single shaft, differential and free turbines to the fixed and rotary wing missions are again reviewed in Task III. Section 5.2 concludes that a free turbine derivative engine using GATE technology is the recommended approach for the rotary wing application. Sec- tion 5.3 presents the engine cost and concludes that the free turbine turbo- prop is too expensive for the General Aviation fleet.

4.5 Fixed Wing Aircraft Point Desi n and Parametric Analysis Beech Aircraft developed airplane concepts that could use GATE-type engines during Task II. The airplane synthesis exercise provided the sizes of engines to concentrate on, and what their benefits would be in terms of air- plane design. This was done for three of the airplane categories defined in Task. I; Category III was divided into pressurized and unpressurized subcate- gories. A pressurized airplane in this class was not in the historical data used in Task I, but it is a type expected to be common by the mid-1980's.

Categories used in Task II were: II -- Single engine 4-place non-pressurized utility airplane IIIU - Single engine 5-6 place non-pressurized utility airplane HIP - Single engine 5-6 place pressurized utility airplane IV - Twin engine 6-place pressurized light executive transport 4.5.1 Computerized Point Design Analysis The main tool used in synthesizing airplane concepts to fit these categories was a Beech in-house computer program designed to match mission requirements and airplane characteristics.

Data inputs to the program included flight performance requirements (cruise speed, take-off distance, range), aerodynamic sanding distance, payload and parameters (drag and lift data), engine characteristics, empirical weight coefficients and empirical landing and take--off factors, as shown in Table XXX. The program iterates the airplane size parameters of power, take--off

s, weight, fuel weight, and loading) until the five above-men-

wing area (wing tioned performance requirements are met. The final size parameters are the program output.

An allowance is made in the fuel calculations for one hour cruise to cover design reserve, take-off, and climb requirements, per prior Beech experience.

An allowance is also made engine for the installed engine weight. This "" weight includes the engine dry weight, controls, exhaust pipe, oil system with cooler, fuel system, propeller and starter generator. Cruise fuel cal- p g p culati.ons are made at a weight reduced from the take-off weight by a value based on past experience. To maintain a safe aircraft design, landing weight is taken mission is to be equal to take-off weight in this program. The assumed, in each case, to be takeoff at maximum gross weight, climb to cruise t altitude, cruise to maximum range, and land. Tradeoffs between reduced pas-- senger and fuel load -vs- extended range or increased speed would be evalu- ated in any follow-on program, to illustrate market features of this tra- deoff. The calculations, however, were made at maximum takeoff weight to illustrate the substantial payoff of GATE turboprop technology at the most adverse mission profile (some current General Aviation aircraft cannot fly quoted range or cruise speed with both maximum payload and a full fuel load).

E A Drag calculations include profile drag, induced drag, and propeller slip-- stream drag. The values of induced (wing efficiency factor) and slipstream drag are inputs: profile drag is calculated in the program, using input values of skin friction coefficients, fuselage equivalent flat plate areas, landing gear equivalent flat plate areas, and other factors.

Lift coefficient values are somewhat higher than most of today`s production airplanes, but are consistent with modern airfoils without unusually complex flap systems (a safety concern for the skill of most pilots in Categories II, IIIU and HIP). Values of were adapted from other Beech lift coefficient proprietary advanced airplane studies.

Baseline engine data were supplied by Teledyne CAE for the C9 engine (Model 2010).

Table XXX Airplane and engine character.isitcs shown in were eventually selected as likely for each category in the mid-1980`s for GATE engined air- planes. These data were the aircraft synthesis program, with the run in results shown in Table XXXI. Compared to current aircraft in each category, the point designs fall in the upper range of gross weight, consistent with their advanced performance capabilities. Because each category represents a range oT marketable aircraft (consistent with the current scenario), it would be expected that availability of GATE engines would also result in a range of derivative aircraft designs of greater and lesser gross weight.

Table XXXI also validates the observation of Task I, that the GATE engines will be flat-rated. Because the engines are sized for power at cruise, the takeoff power required is seen to vary from 88 percent (Category II) to 55 I^

` ^

I L percent (Category IIIP and IV) of maximum sea level takeoff thermodynamic power available. This will ensure engine design margin for both hot day tak- eoff and structural integrity and durability, because none of the engines will have to operate at maximum turbine entry temperature for any significant portion of its life.

i1

The weight calculation equation is based on empirical factors and exponents for turboprop airplanes. Weight output from the program was further analyzed to get a more detailed indication of gross weight and installed engine

I

U

weight.

The program was run a number of times to provide airplane design sensitivity information for Teledyne CAE to use in engine size and characteristic tra- deoffs, leading to optimum engine selection. Typical output data for Cate- gory HIP, are shown in Figures 60 through 65.

These results were used by Beech designers as a basis for airplane layouts; the three-views shown in Figures 66 through 69 resulted. The engine's small size is made apparent in the single engine airplanes by' the rather narrow, pointed nose. Inlet and exhaust pipe sizes are small and unobtrusive. Small engine size is made more apparent by the very small nacelles in the twin engine airplane front view.

l^ A final synthesis was undertaken to relate the (assumed) conservative air- craft improvements to maximum potential aircraft improvements, and the conse- quenL- effect on the engine requirements. In Category IIIP, a 15 percent (additional) cruise drag improvement was assumed, representative of maximum cleanup of the airframe. Payload, range, cruise speed, and takeoff/landing distance were held constant. Results included a 7 percent reduction of tak- eoff power required, and 17 percent reduction of cruise horsepower from 214 to 177 kW (287 to 237 hp); TOGW was reduced 7 percent and empty weight 8 per- cent. Most important, an additional 17 percent reduction of fuel usage was calculated. While these results are deemed achievable, no cost analysis was performed to estimate the manufacturing system changes required to achieve the low-drag airframe.

4.5.2 Aircraft Price Analyses First order approximations of average equipped airframe retail sales price were made for the final four GATE airplane configurations. They include the total airframe, the engine mounts, and propellers and average avionics equip- ment (a value which can range widely) but does not include engine price.

These prices were based on internal, preliminary estimating methods used at Beech, and on the rate results of the Task I market survey. The market sur- vey provided an indication of the number of units over which development costs could be amortized. Pricing policies vary greatly from company to com- i pany in the General Aviation industry, hence the methods used in setting prices are highly proprietary.. The retail prices shown in Table XXXII are probably conservative by industry standards.

Changes in equipped airframe retail price per pound of gross weight change are also tabulated in Table XXXII, to use with the parametric trend curves plotted from results of the airplane synthesis program. These values are i based on airframe and propeller costs only, i.e., without avionics, engines or interior features, This provided a means of changing the price with changes in engine weight or SFC, at a constant mission for comparison of var- ious engine types.

The prices shown are representative of the airplane types shown in the con- ceptual desig n sketches. In Category II, the design concept chosen for the illustration and price estimate corresponds to the most complex airplane in that category.

A lightweight, fixed gear airplane in this size range would have a lower price.

The market survey of Task I was based on historical data that did not include a pressurized airplane in Category III. The first piston engined airplane of this type has recently appeared on the market. It will probably be a common type by the mid-80's. This type airplane can be expected to absorb some sales from the light twin market, as well as generate new sales. With this in mind, the original projected annual sales figure from the market survey for Category IV was consequently reduced from 4400 to 35 0 0 for price estimat- ing purposes.

4.5.3 Noise Analyses Noise for the GATE--engined airplanes was in-house Beec estimated using an proprietary program for the criteria from FAR 36, Appendix F.

A propeller speed of 33.3 RPS ( 2000 RPM) was used with the 2 . 03 m (80 inches) to 2.11 m (83 inches) diameters shown on the aircraft three views. Propeller noise was assumed to be dominant; because of the low , jet velocities of the turboprop, and the buried inlet, no special account was taken of the noise increment added by the engine; this calculation would be addressed in a follow - on prow gram. The results, shown on Figure 70, show that the GATE - powered aircraft will improve the General Aviation fleet noise picture, and are within 1.6 db (worst case, Category IIIP) of the most severe EPA 1985 criteria.

4.6 En ine - Aircraft Synthesis The previous section described the integration of the four point design air- craft with the baseline C9 (Model 2010) turboprop engine.

As a further step toward definition of the optimum engine for each of the aircraft categories, it was necessary to evaluate the effect on aircraft design and performance installing each of the other three candidate engines.

To accomplish this task, the aircraft parametric analyses were reviewed and the data found to be reasonably linear ( Reference Figures 60 through 65) in the area of the design point. It was therefore assumed that the relatively minor i mpact of the engine differences on the overall aircraft design could be addressed in a linearized perturbation analysis. The influence coeffi- dents - for this analysis are summarized in Table XXXIII.

It is evident that increasing mission requirements from Category II through Category IV result in increased sensitivity of the aircraft design to the engine characteristics. Increased engine weight was treated in the analysis as if it were an increase of aircraft payload. The baseline estimated engine performance characteristics at the several mission points were presented in

Figure 71 presents the engine Section 4.4.1: weight and price by configura- tion and production rate respectively, versus sea level maximum thermodynamic power. Sea level maximum thermodynamic power was used as a canvenience for comparison to conventional turbine engine technology, since it i s a more con- f sistent measure of engine capacity than cruise power at varying altitudes and flight speeds.

An iteration procedure was then used to determine the effect of substituting any of the three alternative engines into the baseline aircraft.

Baseline data were perturbed by the increments of engine weight and changes of SFC shown. in Figure 71 and Table XXI respectively using the influence coeffi- cients noted above.

A net change of takeoff gross weight was calculated from the sum of these two effects, and a revised cruise power determined from the J influence coefficients.

This led to a revised value of thermodynamic maximum power at sea level, which in turn leads to a new cruise power, and engine weight increment.

Engine ST'G was adjusted for size effects. The calculation was iterated until it closed upon a final answer of cruise power, from which the other characteristics Caere determined.

Typical examples of the iterations for the effect of different engines on the Category II, IIIP and IV aircraft are Three different shoran in Table XXXIV.

engine configurations for each category are shown; the baseline C9 engine is representative of the Task I data furnished to Beech for the baseline air- craft studies. The adjusted C9 data represents an update of the engine per- formance during Task II, and the AC11.3 perturbations are shown as indicative of the effect of a substantially different engine cycle.

The payload, range, cruise speed and cruise altitude were held constant.

} The baseline Category II aircraft gross weight was 1313 kg (2894 lbs).

With the adjusted C9 engine, the gross weight increased by 2 kg (5 lbs), the fuel eight increased 1 kg (2 lbs) and the aircraft retail price increased $100 from $62,500 to $62,600.

These changes Caere due to an increase in fuel con- sumptioa from 82.4 to 82.8 ug/J (0.488 to 0.490 Ibs/hr/ehp). Installing the -, AC11.3 with a lower fuel consumption of 77.7 ug/J (0.460 lb/hr/ehp) reduced the aircraft gross weight by 17 kg (36 lbs), although the engine weight t) increased by 5 kg. (11 lbs), and the fuel weight was reduced by 11 kg (241 lbs). However, the more expensive engine (AC11.3) increased the aircraft j cost $.2100, or 3.4 percent, from from $62,500 to $64,600.

The calculation was repeated For each of the engines as applicable, and as modified by the scale-down factors previously discussed in Section 4.3.

r 4 At the other end of the spectrum, in Category IV, the baseline gross weight was 3127 kg (6894 lbs). The baseline C9 engine required a cruise rating of 215 kW (288 hp); with the AC11.3 configuration, gross weight was reduced by 324 kg (714 lbs) (primarily due to the greater impact of the 9.5 percent SFC a improvement over the Langer range). In this case, the sales price was reduced, from $233,300 for the baseline -0-powered aircraft to $210,800 for the AC11.3-powered. aircraft..

Following completion of the calculations, a survey indicated that the Base- line aircraft gross weight was not varied 'through the iteration procedure by more than 11 percent in any case. Because of engine price changes (Figure 71) and aircraft price changes due to changes in empty weight, a maximum price change of only 15 percent was noted, validating the linearized pertur- bation analysis. This was accommodated in the life cycle cost analysis des- cribed in Section 4.7. The results of the engine-aircraft syntheses are sum-

marized in Table )MTV for the lowest sales price combinations, in terms of

incremental aircraft gross weight, fuel, load, and power required at the cruise condition. The calculations were limited to 3 categories because they illustrate all of the principles involved in the synthesis (Category IIIV is would not

sufficiently similar in mission and other characteristics that it

add any value pertinent to the objectives of the analysis).

analyses are described in Section

The life cycle cost implications of these 4.7.

4.7 L2 C 2 Analysis - For Selection of Optimum Engines A limited life cycle cost (L2 C2 ) analysis was performed to develop criteria by which optimum engines could be chosen. In conventional military practice, where the aircraft fleet owner is predetermined, a 20 or' 25-year life cycle cost calculation can be made with assurance. In the case of GATE aircraft, however, consultations with Beech Aircraft indicated that the lower gross weight aircraft represented by the GATE market - 907 to 3175 kg (2000 to 7000 lb) TOGW - are normally nor-revenue producing aircraft, purchased in a manner very similar to a car, and held for usually not more than five years, fol- lowed by resale.

For these reasons, it was determined that the life cycle cost analysis would be limited to 5 years, and the acquisition cost approached as a convention- ally financed purchase. Figure 72 summarizes the major features of the L2C2 model and the data sources. Table UXV summarizes the input parameters used for the model, derived for the most part by Beech Market research.

Figure 73 depicts the parametric variables used to evaluate the effect of variable flight time and TBO on the life cycle cost. Again, consulting Beech Market Research indicated that there is no fixed average mission for any of the aircraft in this category; rather a baseline can be assumed, but it is known that various owners will fly the aircraft at widely varying utilization rates per year. Therefore flying hours were also approached parametrically, as were potential fuel costs during the coming decade.

IT are depicted in Figure 74 The results of model calculations for Category for each of the candidate engines (the AAAC--15 engine could not be scaled to the cruise power requirement of Category II as described in section 4.3,1).

It is seen that the nominal 5--year Life Cycle Cost of the aircraft is in the order of $140,000. The CC-20 configuration results in the highest life cycle cost value, but not substantially over the other engine configurations; nevertheless, the basic purchase price of this aircraft would be $10,400 greater than the aircraft powered by the simpler C9 engine, which would miti- gate against a market for this design.

The AC11.3 engine is seen to offer a life cycle cost close to the C9 configu- ration, primarily due to the reduced aircraft price (because of its reduced duced SFC) compensating for the increased ` engine price. However, when the size effects presented Settioq 4.3 are applied to engine performance and ill included in the and life cycle cost analysis, the effect. on S#C and weight.

{ calculated:, a measurable increase 'of cos'b is noted above the C9 configuration h T erefore, the C9 engine, the simplest and lowest cost configuration, is most

applicable the Category e of some doubt on the

to 1l a ircraft... Becaus valid-

ity of the 15 percent. resale value assumed n the arigit model, a separate calculation was performed for a 30 percent resale value; it later deter- was mined tc be more ?realistic. The result reduced the Life cycle cost by some 7 percent, but sloes not change the resulting, engine choice...

Figure 75.illustrates a corresponding calculation for the Category IV air- In craft. this case, all four engine configurations are applicable, and a nominal life cycle cost over a 5-year period is $. 3a0,.00Q. The AC11.3 configuration is seem to be optimum, with the C9 close behind; the other two confLgurati.ons are not marketable, especially when acquisition price differ- ences are considered.

Variations in the life cycle cost with fuel, price and flying hours were also I, calculated; Figure 76 presents the information for the Category II aircraft, and indicates a 39 percent increase of cost as the price of fuel increases from the current level of approximately 18.5 cents per liter (70 cents per `3 gal.) to a potential 53 cents per liter ($2 per gal.) level in the late Si 1980`s. milar data are presented for Category TV aircraft on Figure 77; a 36 percent life cycle cost increase indicated is for the baseline flying hours per year, for the same increase fuel costs.

in The percentage change shown for the calculation would not differ significantly for the AC11.3 engine.

The 9 C 2 analysis, therefore, eliminate& the AAAC-15 and CC-20 configurations from contention. Figures 74 and 75 also show that acquisition cost is, as expected, a major market driver, consisting of between 60 and 75 percent of

Life Cycle Cost for the owner. current prices, fuel represents

At only 15

percent of the life cycle cost; however, , as fuel prices increase per national expectations, fuel could became as much as a 34 percent factor in General Aviation. This places a large emphasis on the need for development of GATE turboprop technology - bath to matte available the multi-fuel capability of the engine (to compensate for the expected unavailability of 100 octane avia- tion gasoline), and to reduce SF'C levels, below current turboprops, to con- serve petroleum.

4.8 GATE Point design Aircraft with Reciprocating Power To obtain an indication of the GATE engine benefits, and to separate the air- craft improvements from those due solely to the engine, estimates were made for Categories Ir And IV (lowest and highest TOGW) of the GATE airplanes with typical: piston engines.

Two comparison studies were made: First, for con- scant aircraft size but variable performance; second, for constant aircraft performance, but allowing a i rcraft size to vary for the design with recipro- cating engines.

Tables XXXVI and XXXVII show the results when the airplane size is held con- ' scant for Categories TI and.

IV, and the mission is allowed to vary, using { equal cruise power for both engine types. The numbers in the left column in each chart are the same as in the GATE-derived designs of Tables XXX and XXXI. The right column shows values for a piston engine version of the air- plane. Changes and percentage changes are shown in the center.

c The following assumptions were made for this comparison: o The take-off weight was held constant.

G o The empty weight was increased by the amount of the increased installed engine weight.

The ' o fuel plus passenger load was decreased by the amount of the increased installed engine weight.

o One-half the empty weight increase was taken from payload.'

o One-half the empty weight increase was taken from fuel.* o Cooling drag was added. This was 10 percent as a representative value.

o For the twin, drag was also increased by 7 percent to allow for lar- ger nacelles. No drag addition was made for the single, because of the nose location - this is an assumption decreases the advan- which tages apparent to the GATE turboprop.

o Recip rocating engine 5FC and weight characteristics were taken from 1977 manufacturer's data, due to the uncertainty of projecting poten- tial advances in SFC, power/weight and the consequent engine price increases. Reference S offers an opinion on these factors.

o Takeoff distance of 2,000 feet or less is judged adequate.

These values were judged worthy of comparison by Beech, since there is no clear figure of merit for the aircraft types under consideration. Other " assumptions could have been used, but would not have changed the main thrust z of the conclusion as to superiority of GATE engines.

i The added weight of the reciprocating engine is seen to reduce range substan- tially (26-34 percent), and cruise speed only slightly (3-6 percent). Tak- eoff and landing distances were found to be equivalent.

^S Another way of comparing GATE and piston-engined airplanes is to hold the mission requirements of speed, range and payload, and let the airplane grow as a result of using an assumed piston engine. The results of this estimate for Categories II and IV are shown in Tables XXXVIII and XXXIX. Take-off distance was allowed to change in this simplified calculation. The engine power requirement increased because of: higher weight and increased drag caused by engine cooling, the larger wing and, for the twin, larger nacelles.

Fuel required goes up in proportion to the power increase, For the Category II single engine airplane, the cooling drag was estimated to require 8 per- cent more power, and the larger wing, an additional 4 percent power. Corres- ..:. .

ponding figures for the Category IV twin are 9 percent and 6 percent, Addi-

tionally the train required 5 percentoe pa g er Iseeause of the increased

t

iiacelie profile drag. Hach of these fact ors is compounded from the assuunp

, tion of equal mission r6quirtfaeht6 the a dded weight and drag of the reci- pr ocating engine onsiderably higher power 'quiremthts, results in c re which i drive the results of the design.

L Therefore, with a constant size airplane; the GATE powered versions show 26-34 percent more range, 3 = 6 percent ihore speed, and 14--22 percent more pay- ,J L ^ Ioad. When the airplanes are compared on the basis of a constant mission

(Table: XL) the GATE powered versa`oils show a reduction in takeoff gross weight

(TOG[) of 12 to 20 percent, a reduction of empty weight (which is propor- tional to airframe price) of 18 to 25 percent, a reduction of cruise power from 20 per c ent and a reduction of fuel. required of 12 to 20 percent.

to distanc e for Category11 shows. a reduction of 13 percent; hour--

eVe r, in Category . 1 V, the takeoff d is tance is in c reased 22 percent which is I'd considered quite acceptable. The method is a simplified approach, using one set of assumptions for a complex coaiparisbn problem: This type of comparison could be done with other, more coinpreheiisive methods; the approach indicates definite and Large advantages for the GATE type engine: These values were converted to dollar increments using the same model., as shown on Figure 78.

The increment is ' - seen to be a 13 percent improvement . fbr Category . 11 air- craft, and a 20 percent ituprov'embAt.of life cycle cost for the Category 111 aircraft: 4.9 Technology Appl.icabil:;tZ and lJo,r.th The objective of this sub-task was to identify large engine and other tech- n.ologies applicable to GATE powerplants, and to assess their value , to the program. Throughout this study, minimO price has been demonstrated as a :1 Major output need o f the study. Three s6 , dkces of price reduction in the 1988 timeframe have been identified: o advanced performance technology (more performance per pound of sim- plified components) o substantially increased production 'rates o improved manufacturing techniques on specific components B This section focuses primarily on materials and manufacturing techniques which support the feasibility of achieving lorry cost on advanced aerodynamic and high temperature, -i.e., high performance components. The technology advancements required in the latter areas was described in Section 4 i and is amplified in Section Ere 4.9.1 Technolog Identification v t Figure 75 sumu,arizes several technologi6 identified from ongoing NASA or DOD-sponsored afforts, most of which involve engines much larger th=n the J proposed GATE engines. Many are nonetheless applicable, if suitably adapted, to such small sized engines. Other technologies incorporated into this sec-- N tion derive from ongoing proprietary work it Teledyne CAE and in the compo- 4o vent fabrication industry. The timeframe of the GATE projections - a readi- development in 1988, i.e., 6- year to into full tress go scale engineering a for feasi- lead time, and a 2 - 4 year .Lead time on availability of materials bility demonstration, forces a degree of pragmatisim into the projections.

The NASA MATE program is providing substantiation data for new materials techniques, but in a radically different environment and shape than GATE.

Therefore, MATE references in the figure imply a similar type of program, rather than a direct technology transfer. From a qualitative point of view, the technologies are described for the C9 engine, but are equally applicable to the AC11.3. They are summarized below: Turbine Nozzles: High temperature materials, possibly coated with the NASA Yttria-stabilized thermal barrier coating. A slight probability exists for the application of ceramics, either in the total nozzle unit, or as a compo- site, e.g., trailing edge. In any case, the objective is to provide a high integrity component (including ba.ckface shroud) with minimum parasitic cool- ing, at a low cost. The ceramic rotor blade shroud and thermal barrier coat- 1 ings effort of MATE could apply directly to the turbine 'inlet nozzle design for GATE, as a static part. Substantial basic data transfer would be expected.

LE

Rotor: As shown in Section 4.4, an integrally cast wheel appears feasible in a conventional alloy; it could be much enhanced as to thermal fatigue by J r' local (tip) directional solidification. Alternately, new developments in rapid solidification, ultra--fine powder metallurgy manufacture also show promise for increased temperature resistance.

The powder metal disk work under MATE addresses axial flow compressor turbine rotors only. Thus, the integral radial turbine GATE rotor represents a dif- ferent shape, with varying material strength requirements between the disk portion and the blade tips. Directional, solidification techniques which are being developed at Teledyne CAE and throughout the industry for separate axial flow blades would probably be adaptable to the integral blade tips on GATE, but represent a significantly different problem statement and research risk as to represent a new, unproven technology.

Abradable Coatings: The abradable centrifugal compressor shroud coatings on MATE do address an applicable temperature range: 700 degrees K (800 degrees F) -- but at far lower tip speeds than the specified 670 m/s (2200 ft/sec); since the abrasion forces are proportional to a power of tip speed, a consid- erably different environment is imposed by the GATE conditions. A similar comp.'rison could be offered for turbine shroud coatings, where GATE requires a 1365 degree K (2000 degree F) plus capability, and more important, a compa- tibili'_y with 747 m/s (2450 ft/sec) tip speeds, thus a completely different shape and thermal distortion environment. GATE research effort could thus draw from MATE, but would have to proceed along lines of different coating composition and application. technique, Emissions: It is expected that all of the information generated on emis- sions-reduction research will be applicable.

Noise: Analysis and treatment of noise will parallel the technology trans- t:.i1k16.; .6 ^LLLV_ Vli Lllf^. L+RV.+ tifiLi WL aVa\( +-V. ./ u4lY V?., 61.1.v L. aa.yV VrL..0 aL6.6 L1 ,_ LLV ongoing development for larg e engines, A maj-or so urce of the cost tec hn ol ogy

will be derived from the automotive field, where mass production -- at a cost

--is ' known" to be committed before 1.983. The automobile fuel control will

sense and activate engine emissions, fuel--air, temperature, valve and injec- and display functions.

tion modulation mechanisms, as well as other monitor The .sensors and computational needs are of comparable complexity to an air-- J craft engine control, hence should form a technology base. Environmental, redundancy.and FAA certification requirements will probably result in design changes, but the capability and worth will be proven by analogy.

U

Gears: The baseline approach to gearbox cost reduction i s selective die casting of housings., with new and novel. application of powder metallurgy, near-shape gears. The necessary research proof will be to demonstrate the structural integrity of the gears under turboprop and helicopter loads and environments, which are more highly loaded than currently proven similar pow- der parts.

Shaft: Pitch and yaw rates on General Aviation engines are not excessive, but could result in a requirement for either very stiff (Large diameter) shafts or increased rotor clearances to accommodate gyroscopic forces. This would drive the design toward advanced technology bearings or lube systems.

An alternative is the high stiffness-low density TiBorSi.c composite shaft

shown on Figure 4.9-1, which allows bearing DN reduction without compromising

the set-up rotor clearances. The high i6dulus--to-density value of the mater- ial will allow the simple, minimum heat rejection shaft system (Figure 79) and offer weight benefits as well.

ij

Other material technologies were assessed, and judged not to be in the stipu- lated timeframe for a commercial engine. For example, titanium aluminide, carbon/carbon composites, ceramic (turbine) rotors and unlubricated bearing systems are known to represent benefits to gas turbine engines; Teledyne CAE has been involved with each in military engine research, and foresees a long--term applicability to GATE-type engines, if they are rigorously pursued.

Therefore, the recommended adaptations and innovations from the quoted cur- rently active technology baselines were chosen to establish a credi:ble.and time-of..-arrival consistent input to the Task IV technology plan.

4.9.2 Technology Worth - Evaluation.

Sectiori 3.6 showed that, unless GATE engines could, be produced for a price competitive with current reciprocating engines, there would be very limited, if any, market penetration. Section 4.6..Rresented.the engine price data, and

"

validated the Task T market scenario. The technology advancements (as dif-

fl

fereritr y_ated from production rate-derived price price improvements) were identified

1..

in Sections 3 and 4.4.

To place a value on these technologies, the L C model was exercised with results as shown, in Figure 80. Perturbing the model plus or minus 5 percent in SFC and plus or minus 4.55 kg (10 lb) in weight from the baselines of Sec- tion 4.7 rebL."._ted in an influence coefficient of each parameter on the life cycle cost. In a Category IV aircraft, it was found that 1 percent in SFC was worth $4890 over 5 years, and 0.907 kg (2 lb.) in engine weight was worth $895. The key GATE technology advances were identified in Section 4 (com- m efficiency), Section 4.3.1 (turbine inlet temperature) and Section 4.4.1. Their worth to a Category IV aircraft was assessed relative to a baseline 1977 engine - a hypothetical, sam e -configuration engine of 9:1 pres- turbine inlet sure ratio and 1311 degree K (1900 degree F) maximum rated period, compared to temperature. A $72,510 benefit accures over the 5-year using typical 1977 production engine technology.

aircraft, with the results sum- Similar values were computed for Category II marized on Figure 81 for the average annual savings for the first five mature fleet years. A fleet-wide savings of $342 million per year is estimated by them 'o combining the calculations for Categories II and IV and extending include Categories IIIU, IIIF, IV and the helicopter fleets. The L C sav- were shown in F i gure $ in s compared to reciprocating e n gine -powered aircraft g g e P P 8 g p 78. Thus a successful GATE program would create a market for new products with inherent economy to the owner.

4.10 Task II Conclusions` The proceeding sections lead to the following conclusions: TT 1. The optimum engines for each aircraft category are:' SEA LEVEL THERMODYNAMIC SFC MAX. POWER CRUISE CATEGORY CONFIGURATION ktd (hp) L&/s--1b/hr/lb II C9 129.8 (174) 82.8 (0.49) IIIU C9 or AC11.3 283.4 (380) 77.7--81.1 (0.46-0.48) IIIF AC11.3 417.6 72.7 {0.435 (560) AC11.3 395.2 (530) 74.4 (0.44) IV 2. Market penetration down to Category IIIU is assured. Sales in Category II are highly probable; the point design aircraft (Figure 66) represents the high-priced end of the broad Category II spectrum (Task I market analysis).

require addi- To expand upon the exact degree of Task II penetration would tional aircraft designs similar to the types currently being sold in Category II, i.e., of lower sophistication level. Engine price bogies (reference Fig- ure 3.6-6) have been approached, but the wide aircraft price and performance at this level spread in this category do not justify a complete penetration of analysis.

3. The challenge presented to Task III is to define a common core to cover a 2:1 horsepower ratio wienout excessive performance sacrifice, and with suffi- powerplant to bring its price into an cient commonality to the helicopter acceptable range (defined by Bell as significantly greater than for the fix- ed--wing aircraft) .

in Task ,,1V are: 4. The key 'technologies to be addressed • High temperature - 1504 t 55 degrees K (2250 t 100 degrees F) uncooled radial turbine and associated abradable housing coatings.

• High pressure ratio (6.6-9:1) backward curved centrifugal compressor and associated shroud abradable coatings.)

• A low cost gearbox.

• A low cost, full authority electronic control system.

5. The benefits of the application of these technologies have been quanti- fied in terms of: sj • More passenger-miles per gallon of fuel.

• More passenger comfort.

• Higher flight sFeed.

• More useful. Load per pound of aircraft structure.

• Loner environmental_ noise than the current fleet, and, if necessary, reduced emissions levels.

The overall results - more aircraft productivity per dollar, at a lower energy consumption level and with a better environmental. compatibility.

SECTION 5.0 TASK III: COMMON CORE CONCEPT EVALUATION.

The individual optimum engines are assessed to define the possibility and utility of a common core across the fleet. The performance, cost, benefits, penalties, and an outline of the methodology of up- and down-rating are des- cribed.

5.1 Common Core Candidates The approaches to the common core are a logical fallout from the Task 11 optimum engine investigation. The C9 and AC11.3 exhibit the lowest cost and good performance levels. The highear pressure ratio configurations, AAF!C15 and CC20, did not show a measurable advantage over the two lower pressure ratio configurations, and are loot considered viable candidates as a common core.

Three basic approaches were evaluated to cover the wide power range (greater than 2:1) required for the general aviation fleet: one Frame size plus shaving, two frame sizes plus shaving and a two-frame family, as summar- ized in Table XL.

One Frame Size plus Shaving: This engine, (Table XL and Figure 82a) has a thermodynamic design and gearbox rating of 365.5 kW (490 hp) and envisions shaving the flowpath to cover the thermodynamic power range from 198 kW (265 hp) to 422 kW (565 hp). The shaving would both increase and decrease the flow channel to cover the full power range. The major penalty of this approach is that the engine weight would remain essentially constant at 93.5 kg (206 lb).

This engine weight would disadvantage the smaller aircraft by increasing the gross weight over what would be realized by a lighter engine.

Two Frame Sizes plus Shaving: This approach would add a smaller (dawn-sca- led) engine to complement the larger 93.5 kg (206 lb) frame. This engine would have a thermodynamic design and gearbox rating of 224 kW (300 hp) and would modify the flow channel to span the thermodynamic power range from 198 kCa (265 hp) to 280 kW (375 hp). The smaller engine size (Figure 82b) and lower weight of 65.8 kg (145 lb) would provide a more desirable powerplant for the smaller category II and III aircraft. The down-scaled engine would utilize the larger engine technology, however, parts commonality would be insignificant, thereby reducing the cost reduction benefits of high produc- tion rates.

Two Frame Family: This approach has the AC11.3 flowpath with a thermodynami design point of 422 kW (565 hp) and a gearbox rating of 410 kW (550 hp). The basic C9 flowpath engine is derived from the AC11.3 by omitting the axial compressor and turbine, and part of the reduction gearing, to provide a smal- ler, lighter engine with a thermodynamic rating of 250 kW (335 hp) and a gearbox rating of 205 .kW (275 hp). The size and weight of these two engines are shown in Figures 82c and 82d. The differences in the mechanical configu- rations are illustrated more clearly in Figure 83. Vae AC11.3 configuration consists of a two stage compressor (axial plus centrifugal), a reverse flow annular combustor and a two stage turbine (radial plus axial). The reduction gear consists of a single stage herringbone mesh followed by a planetar s gearset with two compound idlers. The C9 configuration is achiever, by

remov-

ing the axial compressox and static structure sandwich, the axial flow tur- bine, one half of the herringbone gearset (resulting in a single stage heli- cal mesh) and one of the compound idlers, providing a smaller Lighter engine.

The gearbox sizing criterion was established by the Category IIIU take--off power requirement of 205 kW (275 hp), thus the AC11.3 rating was a fallout.

The performance of the two-frame family is summarizad in Table XLI. The baseline AC11.3 configuration has a thermodynamic rating of 422 kW (565 hp) with a specific fuel. consumption (SFC) of 78.2 ug/J •(0.463 lb/hr-hp). This is a refinement from prior estimates due to the Task III discovery that the turbine inlet temperature could be raised to 1504 degrees K (2250 degrees F).

It was estimated that this engine's centrifugal compressor could be designed to match over a 6.6 to 9:1 pressure ratio range with only a minor (perhaps not any) diffuser change. The estimated performance with the axial compres- axial turbine removed was 250 kW (335 hp) and an SFC of 87.5 tag /J sor and (0.518 lb/hr-hp) at 1389 degrees K (2040 degrees F) turbine inlet tempera- ture. The temperature resulted from an assumption of no' nozzle area change between the two engines. The resulting SFC is only 1.8 percent above the baseline C9 (Reference Section 4.4.1).

A further reduction in power can be achieved without changing the design for

production by using inlet guide vanes (IG17) to reduce the airflow and rematch (closed down turbine inlet nozzle area) to a turbine inlet temperature of 1394 degrees K (2050 degrees F) (Table XLI). This power reduction of 20 per- cent, to 197.7 kW (265 hp) is achieved with only a 5 percent sacrifice in SFC.

5.2 Free Turbine Design For the rotary wing (helicopter) application, the torque/stall characteris- tics of the free turbine are highly desirable. Two free turbine designs were evaluated (Figure 84), a turboprop for fixed wing aircraft, and a direct

drive turboshaft for helicopter applications. These designs use the AC11.3

compressor configuration driven by a single stage radial turbine followed by a two-stage axial flow turbine with power extraction from the rear (exhaust) end of the engine. The accessory drive is located at the front. The differ- ential turbine engine was investigated earlier (Reference Section 4.4.3) as an alternate approach for both fixed wing and rotary wing applications, how- ever, the increased cost and weight relative to the faxed shaft engine ruled it out from further considerations. The result is the free turbine turbo- shaft derivative which is the most acceptable approach to the rotary wing application.

5.3 Cost Analysis The relative cost of design 3013, the AC11.3 common core approach; design -0 3011, a free turbine turboprop and design 3012, a free turbine turboshaft is shown in Table XLIII. The OEM (Original Equipment Manufacturer) costs are summarized in Table XLIV. The methodology used is the same as discussed in Section 3.1.3. The, cost distribution and summary for the AC11.3/C9 common core is presented in Table XLV. Seven of the eight components have been divided into A and B cost units and their engine cost ratio estimated. The A units are common to both the AC11.3 and the C9 derivative, the B units are used only on the AC11.3 design. The combustor is common to both the AC11.3 .

and the C9 and is therefore only an A unit. The K2 column is the fabrication technology factor that is applied for production rates greater than 2000 per year..

The production rates for the A units is 15165 per year, which is the fixed wing aircraft turbine engine market estimated for one engine manufac- turer (Reference Table XIV) plus 35 percent spares. The production rate for the B units is 7525 per year, these are the parts unique to the AC11.3 design ' that power the Category IIIP, IV and the AG (agricultural) fixed wing market.

The engine prices and their relation to the projected fixed wing (by cate- gory) and helicopter market is Table XLVI.

summarized in The number of engines q uoted for the fixed win cat e gories assumes one engine manufacturer f^ q g g g^ will capture only half the aircraft market plus 35 percent equivalent engines .

wa rth of spare parts.

LIU The helicopter market is a small percentage of the total, and assumes that one engine manufacturer will capture the total, plus 35 percent spares. The AC11.3 prices were estimated based on using only a compressor and combustor in common with the AC11.3/C9 engine family.

The C9 derivative of the AC11.3 offers the lowest price (7 percent less) for ru the Category II and IIIU aircraft and is 16 percent lighter than the C9 one frame engine - the next higher up on the cost scale. The category II and IIIU eng ines represent slightly over 50 p ercent of the fixed win market and g^ p g y P g are more sensitive than the other categories to small increases in engine price, therefore, the AC11.3/C9 two--frame family appears to be the promising approach to the common core.

The maximum engine price for market penetration is shown in Figure 85 (based on the ratio of engine price to aircraft price used in Section 3.2.1). The AC11.3/C9 prices (using this ratio) shows a potential market penetration of turbine powered aircraft with a selling price as low as $27,000. The engine price analysis is based on the projected quan- tities shown in Table VIII.

The low cost features of the C9 core engine relative to current technology are summarized in Table XLVII.

Reducing the number of compressor and turbine components reduces the engine cost by 26 percent.

New design concepts for the combustor and control provide another 14 percent, and the improved cyc results in a smaller engine for a further 9.1 percent cost reduction. To be competitive in the general aviation market requires a combination of a low cost design and high production rates. The common core approach to achieving these objectives starts with a simple engine (for minimm.un cost), then adds parts to increase the power output.

The resulting high production rates for the common parts further reduces the engine cost.

SECTION 6.0 TECHNOLOGY PROGWI PLAN i. _.

on exploratory The previous tasks have identified the high priority follow- programs requisite to achievement of GATE turbine engine quality, and the resulting payoff to the General Aviation and helicopter fleets.

The primary objective of this section is to present the overall Technology Development plan which converts the study results into representative hard- ware and addresses the associated risk. Another objective is to provide a sufficient number of management decision milestones to ensure logical cut-off points or alternative directions should the technology achievement be below the level required for the investment strategy.

'.1 Figure 86 presents the major milestone schedule and the key decision points for a 5-year plan, culminating in the delivery of a demonstrator baseline engine for NASA testing. This engine would have a test-defined measure of reliability, making it worthy as a demonstrator of what can be achieved in full-scale engineering development of the GATE technology. In each task, the dashed lanes represent the estimated number of design: and/or test modifica- tions required to achieve program objectives for the subject component.

The program is divided into seven, major line items. It builds from a design definition of the engine through critical component demonstration of the tur- bine, compressor, and gearbox to their integration in a demonstrator engine design.

Because of the breadth of this initial GATE study, and its lack of single-en- gine design, detail, the program begins with a series of visits to potential airframer users of GATE technology. These visits will result in an accumula- tion of mission and power definition data, general information on accessory and installation features, and on flexibilities required of the engine ser- ies. These requirements will be integrated into a mission definition for the baseline engines, against which a specification, sizing, flowpath, structural design, and manufacturing cost analysis may be accomplished.

The major decision milestones are also shown on Figure 86. Near the end of the first year, sufficient engine detail design, cost analysis, and manufac- turing methodology definition will have been achieved to allow the design to pass through its first gate -- is the price of the engine low enough to assure penetration into sufficient new markets to warrant proceeding with hardware demonstration?

In the middle of the third year, sufficient test information will have been collected in radial turbine component development to define its turbine inlet temperature capability. The decision is whether that temperature is suffi- ciently high to warrant proceeding with the program. (At this point in the study, it is estimated that a reduction from 1504 degrees K to 1449 degrees K (2250 degrees F to 2150 degrees F) would not prejudice the results of the program; i.e., even a 56 degrees K (1:00 degrees F) reduction would not be deleterious enough to engine size, weight, cost, or performance to lead to _ cancellation of the program).

Also in the middle of the third year, sufficient compressor testing will have been:undertaken to allow assessment of the worthiness of the compressor, as measured against the original objectives and the requirements of the demons- trator engine.

By the fourth quarter of the second year, the gearbox component development '. will be sufficiently far along to answer whether the powder metallurgy gear approach iS sound. Critical here are the achievement of cost reduction and demonstration of a measure of durability of the process relative to strength requirements for a main engine gearbox. In Task I of the technology plan (Program Control) -» alternate concepts will be made available for substitu- tion should the power metallurgy technique not,prove feasible.

At the beginning; of year 4, the milestone question is - should the demonstra- tor engine be built? The answers will be determined from the accumulation of design and component test data, relative to targets established for a worthy demonstrator engine.

It should be noted that in the first year, for each of Tasks 2 through 5, design effort has been overlapped into earlier task time frames to maintain both program momentum and a reasonable capacity for achieving timely mile- stones. In no case is there any hardware commitment, only design labor.

This approach is recommended so that the` total program can be sustained at a reasonable momentum, achieve a rational end point within five years, and con- tinuity can be maintained by tapping off task results as they can most effec- tively be utilized.

F• ' The 30-month engine demonstrator task is key to the proof of the engine con- cept. Initial efforts are addressed to incorporation of the compressor, tur- bine, combustor, shaft and gearbox component test results into an integrated engine design.

Two serial number engines and the equivalent of 1-1/2 engines in spares will be procured during the fourth year. The first engine will be assembled and instrumented for check run in the first quarter of year 5.

The next three months of engine testing will focus on matching and perfor- mance improvement, using minor modifications to the initial design hardware for increase of power and improvement of SFC levels.

In parallel with Serial Number 1, engine Serial Number 2 will be instrumented for structural evaluation, using high speed slip rings and non-contact tech- niques to measure temperatures and strains. Serial : ry urnber 2 will be run dur- ing the disassembly period of Serial Number 1, so that both performance and structural data can be Integrated into parts modification at suitable times.

Additional testing will be performed during the first three quarters of year 5 to probe the performance potential and structural adequacy of the engine design. This test period will culminate in the running of a 25 hour acceler- ated mission durability test, patterned from current test technique research being done for the USAF and USN ATBGG and JTDE programs. In this type of simulated mission environmental test, the primary damaging events anticipated from assessment and correlation of expected flying data will be used. A structural integrity of the rigorous test will be ponstructed to assess the engine design.

Given success in the 25-hour accelerated endurance, Serial Number I can be refurbished with best available hardware for delivery to NASA at the end of year 5. This engine is expected to be somewhat over the production target on SFC, but probably will achieve design horsepower targets. Its integrity and is adequacy for NASA experimental testing will have been demonstrated by the 25--hour test, hence it will represent a reasonable achievement for the NASA investment in the advanced technology.

This one year test program is a judged value; the program could easily be extended to incorporate two years of testing, with higher end result perfor- ^ nded mance and durability objectives. It is recognized that at some e%b length of the program, it is no longer a NASA technology investment; rather it is a full-scale engineering development program.

r t

SECTION 7.0 CONCLUSIONS AND RECOMMENDATIONS A General Aviation market can exist in the late s for up to 31,500 gas 1980 T turbine engines per year. Approximately 95 percent of that market will be in

fixed wing aircraft of 1040-3040 kg (2300-6700 ' ib) gross wei.ght, requiring

engines of 260-560 horsepower. Five percent will be in helicopters requiring engines of 205-317 kW (275-425 hp). Gas. turbine power is not competitive on a price basis below about 1040 kg (2300 lb) TOGW 194 kTd (260 hp); Larger air- craft requires power levels greater than ' 746 kW (1000 hp) and therefor: bey-- and the study scope.

The applicable gas turbine engines will provide nationally significant pro- ductivity benefits to the General Aviation fleet -- more seat-miles per gallon of fael, 1 lighter aircraft, and greater passenger comfort and safety, They will also meet or exceed anticipated Federal environmental regulations.

o To be salable, the engines must approach the current price ranges of equiva- lent reciprocating engines; benefits to the aircraft design will allow some deviation, but initial purchase price (for the fixed wing aircraft) was veri- fied as a primary objective of any development. The helicopter market can tolerate substantially higher engine prices than fixed wing aircraft.

An roach has been defined to a shaft pp power engine family which will meet the marketplace needs.

It combines advanced technology components with new manufacturing techniques in a high-volume production concept The only forseeable competition will be the improved reciprocating engine, based on its current dominance of this market. , Key factors in this competi- tion will be the ability to improve reciprocating engine power/weight at the same time as SFC, durability, emissions, and noise are improved - without excessive cost increase.

The projected GATE turbine engines will have_;better power/weight and equal or better installed cruise fuel consumption.

nP w The connected shaft:turboprop was found to best address the turbine engine requirements; a free turbine or differentially geared derivative of similar components will suit the torque requirements of helicopters.

A turbofan der- ivative from the baseline turboprop engine would answer the needs of a spe- cialized, high performance market segment.

A family of engines covering the 198-422 kW (265-565 hp) spectrum, with con- LV siderable cost-saving component commonality, was defined. The core engine, applicable to the 198-250 kW (265--335 hp) range, consists of a gearbox, 9:1 pressure ratio, backward curved, single-stage centrifugal compressor; a reverse-flow vaporizer plate combustor; and a 2250 degree F maximum rating, uncooled rotor, radial inflow turbine.

By adding a transonic axial supercharging compressor and an uncooled axial flaw turbine stage, power levels up to 422 kW (565 hp) may be achieved. All rM core components are identical, as are many static housings and most of the gearbox.

The resulting engine will have an 11.3:1 pressure ratio at the same ._I 2250 degrees F maximum temperature rating, and 10 percent better SFC. Compo- nent technology levels are significantly beyond current capability.

Achievement of the component targets requires research to define and attenu- ate the risk of development. The results of a successful program will also t be applicable to cruise missile powerplants and turbogenerator systems.

These combint- to make the results of such a program applicable to small engines across the board, and thus will be worthy of follow-on NASA invest-

ment. }

The study showed a high payoff national, industrial, and environmental impact n~ on General Aviation, and a multi--mission applicability of the advanced compo- nent technology. It is therefore recommended that NASA develop and implement a five-year GATE component and demonstrator engine program.

This study phase reduced the world of possible gas turbine configurations Co ^^ a bounded, focused, multi-purpose family of engines. Because of the study -' breadth, it might be deemed desirable to interpolate a more in-depth design concept validation phase prior to initiation total program.

of the This }phase would be a portion of the 'cask I described in Section 6.0.

_.

CRUISE ALTITUDE 40-1 12- TURBINE

I

ENGINL O DOMINANT O^ L1\ 2U- O W W LU U. W 4 - PISTON ENGINE q PISTON POWER DOMINSrR® O TURBINE POWER REF. AVIATION WEEK & SPACE TECHNOLOGY 3 MARCH 1977 0— 0— a 0 0.5 1.0 CRUISE MACH NO, 32472 Figure 1. Areas of Power Plant Dominance.

I 0 ESTIMATE 1985 NOISE &EMISSIONS HE GS BEECH BELL 5 CATEGORIES OUTLINE DISTRIB, pE NTIFY OF APPLICAS LF INITIAL DRIVER DEFINITION REPROJECT MARKET DOMAINS OF TURBINE POWER.

"'E CONVENTIONAL 1985 COST, RANGE, ALT,. VEL NO SEGME NT S ENGINE MISSIONS & MARKET PROJECTION PAYLOAD COMFORT. SAFETY LCC SUPERIORITY AIRCRAFT CHARACT E R ISTILS WHAT IF TELEDYNE CAE ENGINE TECHNOLOGY PDSTULATION ASSEMBLE PROJ E CTIONS - ALL ENGINE TYPES' WT,. SIZE. PE RF. CO57.

DATA BASE F °3MONTHS ORAL PRESENTATION FIXED, ROTARY WING MARKET FORECAST 1985 AIR VEHICLE • REUUIRED CHARACTERISTICS MISSION PROFILE BY MARKET SEGMENT RANGE OF APPLICABLE TURBINE POWER & SFC REQ'TS NOISE & EMISSIONS REGS, RANGE OF APPLIC. OF OTHER ENGINES SUMMARY OF DRIVERS--REQUIRED ENGINE FEATURES& PRIORITIES TASK 2 APPROACH 25620 Figure 2. GATE Task I, Market Analysis Study Plan.

COMPRESSOR COMBUSTOR

RE[

GEC J4-tuv Figure 3. GATE Turboprop.

F G H HAI (DIA) EXHAUS T .

i IDIA) ACCESSORIES j 25194 Figure 4. Installation - GATE Turboprop.

mr-:F)l ire inns rrnnnDD1:ecn© rnn1P1 ICTnR

! UHb1Nt 31455

Figure 5. GATE Turbofan.

- - -G.

f -r^ C - - -- ^^T D EXHAUST INLET

L__ -J

25196 Figure 5. Installation - GATE Turbofan.

GATE CURRENT TECHNOLOGY 32458 Figure 7. Comparison of Task i Baseline GATE Turboprop With Current Technology.

1.0 CR = 1.0 WT = 227 KG (500 LB) v 08 OE-M PRICE = $59,870 O CR = 0." VVT )+0.24 q 0.6 cc WEIGHT IN KG U) U CR=0.152x10)+0.24 \\ 0.2 WEIGHT IN LB 50 100 150 200 250 KG 100 200 300 400 500 LB Figure 9. CR Formula and WEIGHT Base OEM Pricc Figure 8. CR (Cost Ratio) Relationship.

1.0 )N w U Ir (L > 0.1 g w cc 0.01 100 1000 10000 100000 PRODUCTION RATES (UNITS PER YEAR) 32465 Figure 10. Relative Price Vs. Production Rates.

ENGINEERING AND CAE TURBINE ENG. & AIRCRAFT^— \ ANALYSIS

i r 1988 - 1976

A3RC S CATEGORY CO ' ST PER KILOWATT SELECTION ( HOFlSEPOWER) L._ — I CURRENT MARKET COST PER KILOWATT JHORSEPOWERI ENGINE PRICE ANALYSES I CURRENT MARKET ENG.

PRICE TO AIRCRAFT PRICE RELATIONSHIP It CURRENT MARKET 1 AIRCRAFT' PRICE 8 AIRCRAFC VOLUMFJPRICE —L — --- DEMAND ANAlY51S RELATIONSHIP TURBINE POWER Iggg p15TUR8ED MARKET INFLUENCE MARKEiFDRECAST STUDY IV HISTORICAL 19M UNDISTURBED DEMAND 7REND5 MARKETFORECA5T 32464 Figure 11. Market Analysis Flow Chart.

ENGINE PRICE x100 AIRCRAFT PRICE ^i Q 3600 w a

L

w 2400 n = .

0 `rte e.-i..o^. ^.j....^..^d.r

37000 58000 100000 121000 16000 79000 AVG. EQUIP. PRICE ($ 1 5K GROUPING) 32439 Figure 13. Category I Through III Delivery Volume Vs. Price ($15,000 Grouping).

........

JO 400 J W a ti

z

n -- I- ' ...^..

-50 1280000 560000 920000 1640000 200000 AVG. EQUIP. PRICE ($300K GROUPING) 32442 Figure 14. Category IV Delivery Volume Vs. Price ($300,000 Grouping).

J LU H Z(O 200000 440000 920000 1160000 1400000 680000 AVG. EQUIP. PRICE ($300K GROUPING) 32443 Figure 15. Category V Delivery Volume Vs. Price ($300,000 Grouping).

eliuuu 20000 us LU 16000 > —J w Z 12000

O

BODO 64 68 70 72 74 76 78 80 82 84 Be 88 32445 DATE ANNUAL: 1/63 - 1/88 Figure 16. Delivery Trend and Historical Projection, Categories I Through III.

1988 UNDISTURBED FRAPOLATION 10 YEAR V) Cr

w

w 1600 5 YEAR

D

15 YEAR Z

J

Q

O f — 800 64 66 68 70 72 74 76 78 80 82 84 86 88 DATE ANNUAL: 1/63 - 1/88 32451 Figure 17. Delivery Trend and Historical Projection, Category IV.

1988 UNDISTURBED XTRAPOLATION 15 YEAR 10 YEAR U) F 1600 w

J

U' 1200 5 YEAR

z

J 800

O 0 1 1 1 1 1 1 l l 1 1; I 1 I I I I ^ ^ I I I 1 1 32467 T1 T T 64 66 68 70 7 1 2 74 76 78 80 82 84 86 88 DATE ANNUAL: 1/63 - 1/88 Figure 18. Delivery Trend and Historical Projections, Category V.

24000 6 YEAR 10 YEAR 20000 15 YEAR LU 16000 LU 12000 DATE ANNUAL: 1/63 - 1 /88 32456 Figure 19. Delivery Trend and Historical Projection, Categories I Through V.

iNII IAL ($20,000 TO COST $35,000) OPERATING ($10/HR) ENGINE COST NOISE (85 dBA) (INTERNAL) (73-89 dBA) (r-.XTERNAL) WEIGHT LESS THAN 62 KG (180 LBS) FUEL CONSUMPTION GATE G A E TAKE-OFF RATING TAKE- 261 T ENGINE ENGINE 84 TO 101 ug/J 261 56 KW (0-50 TO 0.60 LB/HR-HP) 350 75 SHP) MPTION PT/J EMISSIONS TIME BETWEEN CONTROL OVERHAULS 3000 TO 5000 HOURS MAINTAINABILF Y 32460 RELIABILITY Figure 20. GATE Requirements for the Light Helicopter Market.

304.8 T HOOD FLYOVER AMERICAN. BEECH. CESSNA, ROCKWELL. PIPER FT1 DATA CORRECTED FOR RIC O 101 90 ry^yr L« D FAR 36/ICAO AFTER 1873 Q -- - — q FAA 198011CAO J W 91 J 80 q O°O O w z D: z :3 a y W V/ DRAFT w D r FAA 0 0 q NPRM 86 0 75 ° 0 z O

- m p — - D EPA 1980/85

0 n. U 81 70 II LEGEND O SINGLE ENGINE TWIN ENGINE q 76 65 3000 5000 7000 9000 10000 11000 12000 0 1000 2000 4000 6000 8000 AIRPLANE MAXIMUM CERTIFICATED TAKEOFF WEIGHT IW1 - POUNDS 0 600 1000 1500 2000 3000 4000 5000 5700 KILOGRAMS 25625 Figure 21. GATE Noise Levels - 1976 GAMA Data.

i PERFORMANCE

- LOW SFC: 76.1-93.0 Ag/J (0.45-0.55 LB/HP-HR)

EQUAL TO PISTON ENGINE HIGH POWER TO WEIGHT - 2 --3.5 X PISTON ENGINE

• INSTALLATION

REDUCED FRONTAL AREA - 1 /2 PISTON ENGINE COOLING DRAG - ZERO DIRECT MOUNTING TO HELICOPTER GEARBOX

• COMFORT

SMOOTH RUNNING - LOW VIBRATION LEVEL- QUIET - SUBMERGED INLET/LOW EXHAUST VELOCITY

O CONTROLS

SINGLE LEVER ELECTRONIC CONTROL

• MULTI-FUEL CAPABILITY

32435 Figure 22. 1988 Desirable Turbine Engine Features,

^

i ELECTRIC STARTER

S OPTIONS

• GENERATOR • HYDRAULIC PUMP • BLEED AIR — PLUG IN AUXILIARY COMPRESSOR CONTROL -- INTEGRATE • ENVIRON!aENTAL BLEED AIR COMPRESSOR/HEAT EXCHANGER/ EXPANSION TURBINE 25747A Figure 23. 1988 GATE Accessories.

FIXED WING

•

Ala

NEGATIVE THRUST ON APPROACH

DECLUTCH/BRAKE * VARIABLE PITCH PROP • HIGH INLET - CLEAN AIR/MIN F.O.D.

* ROTARY WING

• FREE TURBINE OR EQUIVALENT ENGINE TO ROTOR SYSTEM • CLC; .E COUPLE ENGINE/HELICOPTER GEARBOX INTEGRATED TRANSMISSION/ ELIMATE DRIVE SHAFT/SEPARATE MOUNTING • AIR CLEANERS 25739 Figure 24. GATE Interface ConsideratLions.

k„ L1 • SUFFICIENTLY LARGE MARKET DEFINED TO WARRANT NASA INVESTMENT

p

® TECHNOLOGIES REQUIRED TO ENSURE MARKET ARE SIGNIFICANTLY ADVANCED, COMPARED TO 1977 1 NOISE & EMISSIONS ARE VOLATILE LEVERS — EX.CESSIVE REGULATION CAN INHIBIT GENERAL AVIATION MAXIMUM ENGINE COMPONENT/MODULE 1tt1``

U COMMONALITY IN 198-422 KW (265-565 HP) ENGINES

IS ESSENTIAL TO RATE/PRICE/MARKET * EXPECT TO SHOW SUBSTANTIAL GENERAL AVIATION FUEL CONSERVATION 32428A t Figure 25. Task I Market Analysis Conclusions.

L f_7_7 HlMl.C V ! G\,r NULVUT i y OZ) Mr-4,1r. rnvirvt f N Avv 261 t 56 KW (375 t 75 HP) HELICOPTER (BELL-TURBINE) TAKEOFF POWER CAT. IV ? ^j 149 . 283 KW (200 - 380 HP) t^ ..-1 i I III ? > CAT. 213-224 KW (285-300 HP) CAT. II ? ; 112-224 KW (150 . 300 HP) 0 5,000 10.000 20.000 15.000 25,000 30.00) 35,000 32427A 1976 RECIPROCATING ENGINE PRICE - S t Figure 26. Engine Price Bogies •- To Ensure a Marketable GATE Aircraft, Turbine Engines Must Approach Prices Competitive With Reciprocating Engines.

HAAt. S m.

TELEDYNE EDYNE CAR TT7.

t=3 MONTHS PROP FAN MANUFACT COhIPONENT 0FAN bYENDOR a. - v STATE OF LER ADVFROP E . 'DATABANI, 111EART - 1985 FUEL CONTROL PROJECRL MARKET SK II.IENT %E SELECT ENGLM NE FROMTASKI COLIPETTfiVE ASSE6IBLE 1 YELin RANGE. ALT.. THRUST (POWER) p. PARAMETRICSTUDY CONC EPTUAL g3.

-3--1 TRADEOFF OPTIMUM A ENGINE PROD. RATE TEMP» PA. BYPASS ^ ENGINE DATA EVAL. SFC ^-..^ COST LAYOtrFB NOISEd EMISSIONS 13ENEMS TELEDYNE CAEILCNTIFY I APPLY PROJECTED LARGEINRINE7 COMPONENTSICOURG . OtHFR RECOMM.

I =7% MONTHS - DATABANK TECHNOLOGIES v t ORALFRESENTATION POWERS SFC VS. COST TRADES OPRMUM ENGINES - DESCRIPTION AF6?TURES AIRCRAFT LAYOUTS EEhsri. A MA I YS15. CPRE. CURRENT APPLICABILITY OF TECHNOLOGIES SYSTEMS COST ANALYSIS ENVIRONMr2WAL IMPACT 33923 Figure 27. GATE Task II, Trade-off Study Plan.

CAT. 1111? + 1V: 123.5 M/S (240 KTAS) @ 5486 M (18000 FT) f CRUISE TURBINE INLET TEMPERATURE: 1264°K (1800'F) LB HP KW m9 LEA/SEC KG /S HR-LB NS 11 0 0.60 280 1 APABILITY: +72°K (130°F) VARIABLE z O EFFICIENCIES 0.56 O LL.

0.52 z ASSUMED, 1 _ CONSTANT { 0.48 140 EFFICIENCIES X LU LL ASSUMED 3 t CONSTANT 1 220-

\

w

^p^ 4 EFFICIENCY 0.44 VARIABLE EFFICIENCY 0.40 8 10 12 14 16 18 20 8 10 12 14 16 18 20 COMPRESSOR PRESSURE RATIO 32459 Figure 28. GATE Category IIIp and IV: Effect of Variable

Efficiencies on Specific Power, and Specific Fuel

Consumption Vs, Pressure Ratio.

A HELICOPTER (SEA LEVEL): 56.6 WS (110 KTAS) CRUISE TUR13INE INLET TEMPERATURE- 1254'K (1800-F) LB M9 HP KW HR-LB ' ZS TI - 3/ - SEC K - G/S 140- 0.60- 1 CAPA131LITY + 72-K (130-F) B PUM S ICI ASSUMED 220- A VARIABLE RIA 'C'E 6 CONSTANT 0 0.58- 72"' Lr V NC IES EFFICIENCIES E F F S EFFICIENCY tz

0 9

IN 200- U) 120- z 0.52- CONSTANT z L) W 41 PL OA8 - IL 110- 180- ASSUMED C.)

C A P AB' LI 2 CONSTANT C Ty Ez\ FFICIENCIES - VARIABLE D- 0- 0.441 100- U) EFFICIENCY 160-1 21 . 1. I_ I I 0.40 8 10 12 14 16 18 20 8 10 12 14 16 18 20 COMPRESSOR PRESSURE RATIO 32462

Figure 29. GATE Helicopter: Effect of Variable Efficiencies

on Specific Power and Specific Fuel Consumption Vs.

Pressure Ratio.

TURBINE INLET TEMPERATURE 1338 0K (1 950-F)

HELICOPTER, SEA LEVEL: 56.6 M/S (110 KTAS)

z

LU

CATEGORY IV: 136.3 M/S(265 KTAS) @ 5486 M (18000IFT)

G Lj- 11+ LL LU Uj L) 3 z W LL z

CIO

z -j 0 LU D

10 12 14 20

16 18

LL COMPRESSOR PRESSURE RATIO 32448

Figure 30. GATE Typical Cycle Sensitivity Analysis; Fuel

Consumption Sensitivity Vs. Pressure Ratio.

SMALL COMPONENTS: 1.19 KG/S (2.62 LB/SEC) FLOW SIZE MAXIMUM TURBINE INLET TEMPERATURE APPROXIMATELY 167°K (300°F) ABOVE CRUISE PRESSURE RATIO $ 20 EQUIVALENT REDUCTION IF ALL-AXIAL i RADIAL i OF TURBINE TURBINE TURBINE EFFICIENCY ; , 6 BYPASS 6 COOLED BLEED- % RADIAL/AXIAL RADIAL/AXIAL r^ TURBINE RADIA L *-PORTI ON ON (COOLED) UNCOOLED 2 .2 RADIAL AXIAL PORTION & UNCOOLED RADIAL r 1300 1500°K 1400 150C °K 1200 1400 1200 1300 2000 2200 OF 1600 2200°F 1600 1500 1800 2000 CRUISE TURBINE INLET TEMPERATURE 32436

^

Figure 31. Parametric Analysis.

CATEGORY HIP & IV: 123 M/S (240 KTAS) 9 5486 M (18000 F7) MAXIMUM TURBINE INLET TEMPERATURE APPROXIMATELY 167°K (300 .

9 ABOVE CRUISE NP KW mg LB LB/SEC KG/S FIR-LB NS UNCOOLED —""`- UNCOOLED 0.60 17 ---- -COOLED (nEFF'Y. + AIR) 320 ---------COOLED ( AEFF'Y. + AIR) `^ UNCOOLED 300 1 E LIMIT / 3:180- 0.56 ^ _ j PRESSURE U 280 RATIO 15 ^g U 0.52 ,^^r_ 160- a CONSERVATIVEf^' z 14- o^ 240- P a 0.48- 140 3 ^^^ 9 '^`^- ^-^ 9 $^ Cn 5 13 G^ LU 0- 120- 9 0 0 0.44 200 20 180-/

100- 0.40 J

1200 1300 1400 1500 °K 1200 1300 1400 1500 °K 1600 1800 2000 2200 O F 1600 1800 2000 2200°F CRUISE TURBINE INLET TEMPERATURE 32466 Figure 32. Parametric Analysis, mg LS CATEGORY IV: HR-LB NS 138.3 M/S (265 KTAS) @ 5486 M (18000 FT) 0.7r ,— CRUISE @ CORE E .R. = 7:1 T.I.T. = 1200°K (1700"") '18 RANGE OF OPTIMIZED (FAN PR)

p 0.6

-7

TURBOFAN SFC 13:11367"K (2000pF)

a

v

43.2mm (1.7 INCH) ) CASING ILIA.

z

Q J 0.5 Li!

20-300 D LL L) LL U w RANGE OF TURBOPROP TSFC C.

a 0.4 co A TURBOFAN MAY _-.—____ _—_-- _--_-- BE A FALL-OUT DERIVATIVE L 0.3 136.3 M/S (265 KNO' 5485 M (1800 FT) SPECIFIC TURBINE ;E FUEL INSTALLED PERFOR INLET CONSUMPTION TEMPERATURE 0.016M2 0.013 mg STATION LB (30 IN2) (20) „K 0.0113 OR NUMBER HR - LB.

0.0' NS (17.5) LB (2: KN SPEED -100%1 -- - - g 2710 1.5- 2 12.92 0.456 2500 3 13.12 0.463 ^ 4 13.71 0.484 -^-°mss * -0.0497 1278 2300 5 16.35 0.577 1.0 4 u 200 - r to `t 6 36.29 1.281 1167 2100 [^.

g -- 1000 100 015- O.

pE 0 0J 833 1500 ).013 0.019 SPEED-95% (20) (30) 1.5 ! .

1506 2710 300 ^- 1389 2500 1 12,78 0.451 ^ g 2 1278 2 12.64 0.446 co 1.0

3 12.55 0.443 1

o^ 1167 2100 200 A9 4 12.49 0.441 s g^ 5 12.95 0.457 1000 1800 r 6 14.87 0.525 100 0,5- 833 1500 SPEED-90% 1.5 0.( 0.0097 0.019 (15) (30) (^ 1506 2710 RS A 1 13.63 0.481 1.0 1389 2500 13.37 0,472 Z1^ 200 u 2 --- --- 3 cl) Gx 17.008 1167 2100 13.29 0.469 12,5) 1000 1800 5 13.57 0.479_ 0.5- Oo 0.539 6 15.27 g6 ^f 0 0 1.0 1.1 JET NOZZLE Figure 34. Engine Turbine Ratio f WITH ACCESSORY AND BLEED LOSSES TURBINE: FUEL INLET FLOW KW TEMPERATURE HP 400 THRUST 0312 °K OR G/S I.B/HR 0.356 KN (70) (80 LS) j 1506 2710 31.0 246 NOZZLE 1389 2500 26.8

f 213

PRESSURE I

Ct 0..26 RATIO LU 1278 2300 22.9 182 1.3 a 300 1167 2100 19.2 152 /1.2 20.4 0.222 f 1.1E .^ (50) 1056 1900 15.5 123 1.10 944 1700 113 1.05 100 f 833 1500 8.6 68 32417 ►

o [_ 0 1 I

^i 0 006 0.008 0.010 0.012 0.014 M2 .I ..1 1 I 10 12 14 18 20 16 22 IN2 JET NOZZLE AREA Figure 35. Engine Shaft Power and Thrust Vs. let Nozzle Area and Turbine Inlet Temperature at S.ea Level Static, 9:1 Pressure Ratio Engine Computer Model.

Z i`- ig RECOMMENDED MAXIMUM CRUISE SCALING POINT /HR HP y g/J - EL LB 136.3 M/S (265 KNOTS) 5486 M (18000 FT) 0.60 Z) 1o0 LB I^^ 128.6 MIS (250 KNOTS) 9144 M (30000 FT) - ° - • - - 0 56 r 40 L 2 0

_

0.52 0 0 U 0.48 80- ^.

0.44 Z >

5 0.40

100 150 250 300 0 50 200 I 0 100 200 400 EQUIVALENT POWER 32471 Figure 36. C9 Engine Performance: Category IV Baseline.

198x LARGE ENGINES POLYT ODIC TODAY EFFICIEuGY T O L/STA TIC TA PERCENT as TELEDYNE CAE 1985 PR OJECTION - -- SMALL ENGINES CC SYMBOLS C CC AC 1978 STATE OF"THE ART ADJUSTED TO AAAC AAAC r_ 1.19 kg/S (2.62 LB/SEC) S2 4 b t5 lU 7z i4 . 1b 3tS 32429 COMPRESSOR PRESSURE RATIO L7 Figure 37. 1985 Compressor Component State-of-the-Art Useable Eff'ic'iency.

RADIUS 180— 7 a Lea b rev/s 1160.3 CORRECTED (RPM) (&3900) SPEED CORRECTED KG/S 1.19 FLOW (LB/SEC) (2.62) PRESSURE ...

9.0 ,RATIO 20 VARIABLE INLET ADIABATIC PERCENT 81.5 GUIDE VANES EFFICIENCY w 80 z r 1 ^ '0 1 — STAGE 1 2 OVERALL CORRECTED rev/s 1367 . 0 1346.2 136-".0 SPEED (RPM) (82550) (75070) (63250) CORRECTED KG/,S 1.18 0.77 1.19 FLOW (LB/SEC) (2,02) 41.698) (2.62) RADIUS PRESSURE 7 180 Y.

1.74 6.65 11.5 RATIO ADIABATIC 160 PERCENT 822 84.4 81.71 EFFICIENCY 1' 1 STAGE 1 THRU 3 4 OVERALL CORRECTED reels 1367.0 1005.6 1367.0 ^1 SPEED (RPM) (82350) (63590) (82350 ► CORRECTED KG/S 1.19 0.334 1.19 RADIUS FLOW (LB/SEC ► (2.62) (0.736 ► (2.82) 7 180 PRESSURE 4,66 3 . 24 15.0 i RATIO ADIABATIC PERCENT 805 83.8 78.2 EFFICIENCY ., N "k LU LU 4LU1C0 F. 4 Ll i 0 01 0 20 40 60 80 100 120 140 160 180 200 AXIAL MILLIMETERS LENGTH 0 1 2 3 4 5 6 7 8 INCHES 32479 _i Figure 40. GATE AAAC 15 Compressor Design Study.

_ t 1 2 OVMALL STAGE CORRECTED rev/s 1246.3 86910 1246.3 (7W80) (52350) t7600) SPEED (RPM) RADI 1.19 0.19 1.19 CORRECTED KGJS FLOW (LB/S) (2.62) (0.42) (2,62) PRESSURE 9.0 2222 19-89 RATIO

AOIAeATIC 1 pEgCENT 713.8

80.9 76.6 EFFICIENCY 4- 3- 2- 60 -90 100 120 140 160 180 AXIAL 0 20 40 MILLIMETERS LENGTH 0 . 1 T 2 3 4 5 6 INCHES 32475 Figure 41. GATE CC 20 Compressor Design Study.

46 .6% E

-a.- -0- --P- -r- % 16%COOLING v 25.6% fla = 0.067 JLLIL- TDISCHARGE = 1254K (I 800-F) fla = 0.021 99.1% 22.0%

Ell

-1-N TIN SD0.5*K f/a = 0.067 167% (442*F) 9% COOLING I 5% r1,-r1,-r>-(1,- 0.5% \30.7%L 0.4% 32453 Figure 42. Combustor Flowpath and Flow Distribution.

GATE STARTING

ENGINE DATA

RANGE /a) = 0.020

x

GATE STEADY STATE RANGE U z ul U_

W = COMBUSTOR

uU

AIRFLOW

O

V = COMBUSTION

E-

VOLUME

m P = COMBUSTOR

INLET PRESSURE

c,

t — —f

KG/S

I

0 50 100

150 200 M8 ATM2

I LB/SEC

0 2 4 6 8

10 12

F-3 ATM2

W

AERODYNAMIC LOADING = Vp2

32470

Figure 43. Combustor Feasibility Analysis.

IU M/S FT/SEC 900 UPPER LIMIT w w LOWER LIMIT M Uj wQ r" 2000 U/co = 0.69 16 06 J B/LB J/9 .3 50 - 120 40 100

F

L

8 0

U

cr

f

° ( 20 40

O F OK 2400- m w Q - UPPER LIMIT 23x0

r

i5 oa ow ¢ -- 1450 M LOWER LIMIT

210 0

w

^. r

1400-

D

cc F' 1900 1300 .

2 3 4 8 5 6 7 PRESSURE RATIO: TOTAL-TO-TOTAL 32473 Figure 44. Maximum Temperature Capabilities of Uncooled Radial Turbine Rotors.

t W Lro-/b = 0.30 Kg/S (0.66 L13/SEC) RADIUS

= 1604°K (2710°R)

T

P = 879.8 KPa (127.6 psia) 2g.Jo ° H 2.18 U2 f

Tr = 1257°K (2266°R)

4 Ut = 759 m/s (2490 FUSEL)

n = 87.9%

CC 80

T = 995°K (1792°R)

P = 109.7 KPa (15.91 psia) L

AH = 627.9 J/g VX

I (270 B /LB) VCR - 0.5

2 - J

ASTRESS = 400 MPa 1 (58000 PSI)

0-- 1 C -f - ---

4Q 60 0 20 80 100 AXIAL MILLIMETERS LENGTH 4 32433A 0 1 2 3 INCHES Fig--e 45. C9 Turbine Design Study.

'^ RADIUS

W 4 - 013 0.2343 Kg/S (0.52 L[3/SEC)

2g

=1 O ° H

= 1.ss

A H = 206.7 J/g (88.9 B/LB)

80 E U

t

= 693.1 m/s (2274 FT/SEC) 17 = 88.0

A - 451.6 J/g (194.2 B/LB)

^

w 60

^

n=87.0 ^ U 2

V/VCR = 0.55

?

V/VCR = 0.4

-4 40

0' 0-^--^-- - -T- —^-- - - -^- --^- 0 20 40 60 80 100 120 140 160 180 AXIAL MILLIMETERS LENGTH F ^ 4 5 7 32434A 0 1 2 3 6 INCHES Figure 46, AC11.3 Turbine Design Study .^ F ---------- W 0. 113 Kg/ ,S (0.39 LB/SEC)

2g Jo A

RADIUS 1.88 A H 139 -6 J/9 (60 -0 B/ U2 4 100-1 AH = 139.7 (60.

U 680.9 m/s (2234 FT/SEC) 03 80- = 86.1 cc ui 436 J/g (187.5 B/LB) AH 60- 86.9 — V/V 0.61 2- r CR — — 40- VNcf3 0 4 +8.20 C) i v 20 0-1 0 20 40 60 80 100 120 140 160 180 200 AXIAL MILLIMETERS LENGTH

0 2

3 4 5 7

6 a

INCHES 32468 A = 1.88 - 147.2 JJg ( 63.3 B /L8) 147.2 J/g (83.3 B /LB) AH 100-- U2 188.1 1 7 - °0 86.8 U = 694 . 9 m/s (2041 Ff/SECS 80- SPEED -1367 rev/s (82350 RPM) V/VCR = 0.48 w 60- .^ 0,8905 2 .

g 22V 238 23V 33B SPEED - 1248 rev/s 40- V/VCR = 0.4 (57760 RPM) 1 SPEED - 963 rev/s (57760 RPM) 20 A H - 363.9 J/g (156.5 B/LB) 0 20 40 60 100 140 160 180 80 120 200 AXIAL MILLIMETERS LENGTH 0 1 2 3 4 6 7 8 INCHES 32449 A Figure 49. AC11.3 Free Turbine Design Study.

K GAS TEMPERATURE 1450°K (2150`F) r.rWgTnrnir 1ArnRu w cc h

w

h 1920 ^ U

-86

J

Z L!!

L!1

U

Z

u,l !11 Er U O 1840 78 U.

w h cc

1 1800_ 4 260

1760 -J 600 650 700 750 800 850 900 WS 2000 2200 2400 2600 2800 FT/SEC TIP SPEED 32452 A

FiaurP 50. Rn'tnr RP1ativP T P.mnAra'tIIrP an d FffiriQnry

^

EFFICIENCY EFFICIENCY (PERCENT) (PERCENT) TOTAL TO TOTAL TO SCALE & CLEARANCE STATIC TOTAL SCALE ONLY RADIAL-AXIAL TURBiNE AC 113 C9 sc 8 0 AAAC: 15 SIZE LIMIT 10.2 mm IN (0.4 3 40 6 8 100 850.9 (33.5) -- — °.

340.4 1.9 3.4) .2) ( 1 DIA, A.

457.2 (18- 0) DIA.

CONFIGURATION C9 CONFIGURATION AC 11.3 DESIGN 2010 DESIGN 3010 ------ 972.8 (38.3)-- -----^^ -- -- 995.7 (39.2) 1—.1 147.: --- 322.6 (5.8) - — (12.7) {16.7) DIA.

DIA.

.rte+ __J 457.2 ^ 57 (18.0) DIA.

DIA.

DIA.

32064A CONFIGURATION AAAC 15 CONFIGURATION CC 20 DESIGN 4010 DESIGN 5010 Figure 52. Basic Engine Configurations: Equal Power - 365.5 KW (490 HP) at Sea Level Static Conditions: Dimensions - mm (in.)

li J

innnn um b ^, A r,.

..

78946 Figure 53. Design 2010 Engine Layout.

so ir-'.,...: :.^.._ ..^..—.-..,---„_..^`-^v—.—.^.J.^..-•---- .r,....,. :.

-^ a ....

': ..._JL'_;.........--..`__ _...____.^..-,.....^._..

a F E

D

AA __r f G H -(DIA)(DIA)

` K

V I.

(DA)

32444 Figure 54. Installation - GATE Turboprop-.

1400 1S00 1800 (7) 9000 1100 1200 9300 °K RADIUS I I (IN) mm CRUISE l ^^`; 4 100 MAXIMUM Go — 2 2 Is r 40 Y 20 0 0 -- BLADE METAL TEMPERATURE BLADE BLADE ROTOR GEOMETRY GEOMETRY ELEVATION ATR=31 ATR-16

JH

STRESS RUPTURE LIFE (HOURS) ATR i6 ATR 1 31 3000 10,000 PLUS CRUISE MAXIMUM TURBINE INLET TEMPERATURE: 15WIC (2250°F) 100 70 300 MAXIMUM MINUS 28°K (SWF) MATERIAL: 1978 EOUTAXED IN-100 32432A Figure 55. Turbine Rotor Design Summary.

4.1 2.5 Sms 2.5 1.5 2.0

7 7

UR

Z7

RELATIVE (310 1.5 COST :S B 10 12 14 is is 20 CPR 32478 gure 56. Relative Cost and Power to Weight Ratio With Increasing Cycle Pressure Ratio (CPR).

10470 1 UM51mr- FUEL CONSUMPTION INLET 96 TEMPERATURE ORC LOCKUP so so 137-

0 so 60 so 4

10 70 so 0, 32481 OUTPUT SPEED - % OF DESIGN Figure 57. Power Characteristics of a Differential Turbine Engine.

D4 — (73.6) (1 4.6) DIA DIA T COMBUSTOR ORC TURBINE COMPRESSOR POWER OUTPUT ................ 365.5 KW (490 HP) FUEL CONSUMPTION.— 89.6A9/J (0.53 LB/HP-HR) WEIGH'? ........................... 918 KG (260 LB) DIMENSIONS mm (IN) 3242.

Figure 58. Differential Turboprop Design 2011.

850:9 (33.5) 345.4 (13.6) DIA POWER OUTPUT ................ 365.5 KW (490 HP) FUEL CONSUMPTION .... 89.6 A91J (0.53 LB/HP-HR) WEIGHT ......................... 111.2 KG (245 LB) DIMENSIONS mm (IN) 32426 Figure 59. Differential TUrboshaft Design 2012.

TUR130PROP RANGE =1 667 KM (900 NCI) PAYLOAD = 453.6 KG (1000 LBS) TAKEOFF DISTANCE = SUM m (2000 Fr) KG LANDING DISTANCE •= 457.2 m (1500 Fly (1000'S) LB (1000,5) S BASELINE • 3 = 4 ^C CD w FUELS F-2 m2 } LU

200—

10— KW HP 600— i Ll

T,O. POWER

O

n

200—

R

r CRUISEpOW ^ 0 0^

90 100 110 120

130 140 150 m/s

180 200 220 240 260 KTAS1 VELOCITY AT 5486 m (18000 FT) 32431 Figure 60, Airplane Size Vs. Cruise Velocity - Category 111P.

TURBOPROP V = 108 m/s (210 KTAS) @ 8486 m (18000 FO PAYLOAD = 453.6 KG (1000 LBS) LB TAKEOFF DISTANCE = 609.6 m (2000 Fr) (1000-S) LANDING DISTANCE = 487.2 m (1500 F') KG (1000'S) [r i r w [ } 2 --)0 KM - - - - I !

I f I ,^„ t ^^; 400 600 800 1000 1200 1400 NM RANGE 32438 ure 61. Airplane Size Vs. Range - Category HIP.

TURBOPROP V = 108 m/s (21 0 KTAS) @ 5486 m (18000 FF) RANGE = 1667 KM (900 NM) LB TAKEOFF DISTANCE = 609.6 m (2000 FT) (1000,S) LANDING DISTANCE = 457.2 m (1500 FT) KG 8 - (10 00 ,S) BASELINE 3- 2- 4 - ul 2 - FUEL WT T- 2 F m 400 - 30- (D < 20- W 200 < 10- HP KW 800 - 600- 400- POWER CL 200 - =:==4^ CRUISE POWER ^ 0 J 0 200 300 400 500 600 700 Ku

400 600 800 1000 1200 1400

PAYLOAD

32437

Airplane Size Vs. Payload Category HIP.

Figure 62.

TURBOPROP FT) V = 108 m/s (21 0 KTAS) @ 5486 m (18000 LB (900 NM) RANGE =1667 KM (1000,S) PAYLOAD = 453.6 KG (1000 LBS) LANDING DISTANCE = 457.2 m (1500 - 8 "1 KG (1000,S) BASELINE

y 2

CD LU

z

2 .4

F T2 M 2 40 0- Z W 200 — --------- 4^^ HP KW ^O .

_), / w 0 400 10 4^10 CRUISE POWER 200 300 400 500 600 700 800 900 m 800 1200 1600 2000 2400 2800 FT TAKEOFF DISTANCE 32424 Figure 63. Airplane Size Vs. Take-Off Distance Category HIP.

i TURBOPROP V = 108 m/s (214 KTAS) @ 5486 m (18000 FT) RANGE = 1667 KM (900 NM) LB PAYLOAD = 453.6 KG (1000 LBS) (1000'S) TAKEOFF DISTANCE = 609.6 m (2000 FT) '. r 8— KG (10 6— ` M rR 1 / 4 FT2 400 m 1^ KW HP 800 600 O SOW eR 4ao Q 200— CRUISE POWER a 0 500 700 80a M 200 300 400 600 1200 2000 2400 FT 800 1600 LANDING DISTANCE 32440 Air ,^nc Size n g Distance - Category HIP.

Figure 64, Vs. Landi 9 9 Y P ;" 8 8 TU RBOPROP 108 m/s (210 KTAS) @ 5486 m (18000 Fi7 V R R A -= 1667 KM (900 NM) AYLOAD = 453.6 KG (1000 LBS) P Uj KEOFF DISTANCE = 609.6 m (2000 FT) TA LB LANDING DISTANCE = 457'.2 m (1500 FT) (1000,5) K (10 a a,S) BASELINE X T.O. WT.

2 1 IS L FUEL 119T.

F-2

m2

3 0

^¢

2 C '^J ? 200 HP K W 60( 44C w CL 20( Mrt C 60 80 100 A 9/1 0.2 0.3 0.5 0.6 LB/HR-EHP 0.4 CRUISE FUEL CONSUMPTION Figure 55. Airplane Size Vs. Cruise Fuel Con: 12.87 M (4221 F1) M F] )9A Figure 68. GATE Turboprop Powered Aircraft Three Views - Category HIP.

I ^ 9.66 M 0.305 M (12 INCH) (31.7 F1) 2.08 M ( 82 INCH) DIA g 2.65 M 3 LADE O O (8.7 Fly GROUND LINE ^.

a .^ / 13.46 M 32411 GATE Twin Turboprop Powered Aircraft Three Views - Figure 69.

9 P Category IV.

101 90 96 85 FAR ^6110A0 AFTER 1973 d a n b FAA 198011CAD ui 91 w 80 ur a a ¢ z ^ ° o 0.

^ q n as 75 a z q 0 n V3 EPA 1980185 q r 81 70 Q LEGEND SINGLE ENGINE O d TWIN ENGINE 76 65 0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000 11000 12000 AIRPLANE MAXIMUM CERTIFICATED TAKEOFF WEIGHT (W) POUNDS 0 600 4000 1500 2000 3000 4000 500G 5700 KILOGRAMS 25625 ^r F ^ 09 ENGINE RETAIL PRICE 800001

0 1)

C^ 60000' C9 1.0 AC 11.3 1.346 5 40000 AAAC 15 1.932 CC 20 1.676 '100 30000 C9 1.0 20000 AC 11.3 1.285 2°pb AAAC 15 1.777 .:s CC 20 1.581 gyp° 0 0p ° 10000 A00°° ,s LB KG ^-- CC 20 SCALING = 300 LIMIT C'} 200— 1 2 3 4 5 6 KW(100'S) 1 2 3 4 6 8 HP(100'S) POWER (SLS, THERMODYNAMIC) 32418 Figure 71.

Engine Weight and Retail Price Vs. Power for Different Production Rates.

0, 0 NON-REVENUE AIRCRAFT: 5-YEAR OWNERSHIP, 20% RESALE

0 L 2 C 2 -FIXED + DIRECT (VARIABLE)ICOST OF

OWNERSHIP - FY 77 DOLLARS

i

MISSION: PAYLOAD & RANGE AS IN AIRCRAFT SYNTHESIS + PARAMEmic HRS/YR 0.

FIXED VARIABLE AIRCRAFT . ENGINE AIRCRAFT + PROP ENGINE

ACQUISITION O) ACQUISITIONM --- FUELM

SERVICING/PREFLIGHT SERVICING/PRE2VGHT

INTEREST INTERESTWN/ MTCE & REPAIR 3 MICE & REPAIR 2

Tt

P

INSURANCE (3} INSURANCEP OVERHAUL OVERHAULP

ij _- HANGAR & MISCO) MISC (3) -

(1) BEECH ANALYSIS-AIRCRAF T SYNTHESIS RESULTS

32447 L.1 (2) TELEDYNE CAE ANALYSIS-ENGINE STUDY RESULTS (3) STANDARD FORMULATION

Figure 72. Limited Life Cycle ( 2 )

L 2 C Features and Data Sources.

FUEL COST: 18.50 TO 530 PER LITER

(700 TO $2.00 PER GALLON)

"AVERAGE" CATEGORY HRS/YEAR RANGE MISSION USAGE: 11 & I I I U 500 1 00 - I OGO

fli p & IV

400 100- 800 32457 T.B.O.: TURBINE - 3500 HRS.

E RECIP. - 1250 HRS.

Figure 73. Parametrics for L 2C 2 ?analysis.

r L 2 C 2 .

$ (1000'S) FLYING TIME 500 HRS/YR 20D ENGINE CONFIGURATION CC20 C9 AC11.3 DOWN-SCALE -- / LOSSES AAAC15 NOT SCALABLE TO CATEGORY II POWER LEVEL D.O.C.

SO FUEL 01 O 3211 OA (1) COST. 18.50 PER LITER (700 PER GALLON) Figure 74. Category II, 5 Year L 2 C 2 Summary. The Simplest Engine (C9) is Best.

L2C2 $ 0OOO'S) FLYING, TIME 440 HRS/YR ENGINE CONFIGURATION CC20 40D AAAG15 C9 AC11.3 - 30O D.O.C.

10a FUEL(1}

L

O

{il COST. 18.5¢ PER LITER 1700 PER GALLON) 32112A

Figure 75. Category IV, 5 Year L LCZ Summary. The Simplest Engine (AC 11.3) is Best.

CATEGORY 11 - C9 BASELINE ENGINE uw ---FLYING HRSJY L2C2 (XG 1000" 200- 30% .110 ioo 0.20 0.30 0.40 0:50 MITER FUEL COST 0.70 1.00 1.50 2,00 MALLON FUEL COST 32477

2 2

Figure 76. L C

Vs. Fuel Cost - Category II.

L'W (XOOO) CATEGORY IV - C9 BASELINE ENGINE S 700 36% 0.50 $/LITER 010 0.30 0.40 k 1 0.70 1.00 1.50 2,00 a, GALLOP!

32476 FUEL COST

Figure 77. L2C2 Vs. Fuel Cost

Category IV.

BASELINE C9 ENGINE CONSTANT MISSION AIRCRAF T DESIGNS acly CAT. IV RECIP 20% GATE .

aoD &L'C' FROM: LOWER EMPTY WT.

LESS FUEL LONGER T90 CAT. II RECIA 13% GATE 0.20 0.30 0.40 0.50 $&[TER 0.70 1.OD 1150 2.0C S/GALLON 32108 FUEL COST Figure 78. L C Vs. Fuel Cost for GATE and Recip. Engines.

THERMAL BARRIER ( NASA), HI TEMP MATERIALS &IOR CERAMICS ` +/ " ` BLARING WO (Ti BORSIC^ SHAFT ` MATE)M.OISC N Lj t ^ F r EMISSIONS a i REDUCTION 0, f^^ ABRADA9LE COATINGS - (MATE) 00!

tir 1ELECTRONIC FUEL CONTROL 27880 GENERAL P .M. TECHNOLOGY NOISE REDUCTION ANALYSIS & TREATMENT Figure 79. Applicability of Large and Other Engine Technologies to GATE.

I^1 L 2 C 2 C.91 KG/s LB) 5-YEAR VALUE OF 196 SFC ENG. WC. (a) CAT. IV .(b) $9890 $$895 A WEIGHT $ 1,202 TECHNOLOGY ADVANCE EFFECT ASFC-°/6 KG (LB) SAVINGS PER A/C 9:1 P.R. COMPRESSOR - I-2°6 A^^ —2 — 2.7 (-6) 15,150 —12 35,300.

UNCOOLED RADIAL. TURBINE -167 ° K T.I.T. —2 . 9 (--28.5) (-1-390°F T.I.T.)

I%AnG —1.1(-2.5) 6,$40 ABRADABLE COATINGS + —0 . 9 +1 % AT) r —2.1 —2.0(-4.5) 14,300 —0.6 (-•-1.4).

TIBORSIC SHAFT --0.61 KC 0 1,250 (-1.4 LBr GEARBOX COST —10% - - 435 3,200 ELECTRONIC FUEL CONTROL - - - TOTAL 76,275 (a) PER ENGINE 28888A (b) TWIN ENGINE Figure 80. Typical Worth of Technology Over a Five Year Period.

AVERAGE OF FIRM 5 114ATURE YEARS $SAVINGS - MILLIONS CAT. II: 9600 AIRCRAFT/YR. = 72.2 CAT. Ill: 4400 AIRCRAFT/YR. -- 191.4 263.6 ADDING CAT. IIIU + IIIP + VI(AG) + HELICOPTER WOUL D INCREASE THE SAVINGS BY 30% CONCLUSION: PAYOFF IS WELL WORTH NASA INVESTMENT zaeas Figure 81. GATE Fleet Yearly Cost Savings From Turbine Engine Technology PAvances.

_ 970.3 (38.2) 129.5 . ^..

it I 315.0 (13.4) (12 .4) IA ."A

Ii- ^L % -

it 416.5 457.2 — (164) (1910) MIA DIA DESIGN POWER - 365.5 KW (490 HP) DESIGN POWER - 422 I<W (565 HP) WEIGHT - 93.5 KG (206 LB) WEIGHT - 92.2 KG (203 LB) (a) C9: FULL SIZE FRAME (c) AC 11.3 COMMON CORE 744.2 It 863.5 (34.0) 129.0 129.5 (5.i)

^5

- (11.2) -- -- (2.4) 40'11,3 416.6 'W - ^ — (15.8) - DIA ( DIA ) DESIGN POWER - 224 KW ( 300 HP) DESIGN POWER - 254 KW (335 HP) WEIGHT - 65.8 KG (145 LB) WEIGHT - 78.1 KG (172 LB) (6) C9: DOWNSCALED FRAME (d) C9 DERIVATIVE FROM AC 11.3 32422' DIMENSIONS: mm(IN) ;i .a Figure 82. Common Core Layouts.

1.11 CONFIGURATION AC 11.3 422 KW (565 HP) 92.2 KG (203 LB) t 3 WITHOUT SHADED COMPONENTS RESULTS: CONFIGURATION C9 r rf TURBOPROP. DESIGN 3011 POWER: 365.5 KW (490 HP) WEIGHT: 100.8 KG (222 LB) I< ^.1 a T URBOSHAFT: DESIGN 3012 J POWER: 365.5 KW (490 HP) 32556 WEIGHT: 73.5 KG (162 LB) ^ t Figure 84. Free Turbine Designs. ., ENGINE PRICE (RETAIL) $ (1000'S) MAXIMUM MARKET CATEGORY IV PENETRATION TIT r li I AC11.3 O Cg COMMON CORE (AC11.3 DERIVATIVE) 2 L- 20 40 60 80100 200 10 3. AIRCRAFT PRICE - $ (1000'S) 294ns TASK TITLE YEAR I YEAR 2 YEAR 3 YEAR 4 YEAR 5 I PF,OGRAM CONTROL - DESIGN & COST MGT.

- ENGINE SPEC. & LAYOUT -MANUFACTURING STUDY & DTC ^ IS PRICE C OMPET$TIV ?

2 RADIAL TURBINE COMPONENT REVEL. - DES. & FAB.

- MATERIAL, COATIkG, FAB.METHODS RESEARCH T IS T.I T CAPABIt_Ii -AEROMECH. TESTS: SPIN, PRESS., TEMP.

IS I f HIGH EN UGH?

3 COMPRESSOR COMPON . DEVEf..- DESIGN &PROCURE ^ _ COMPWORTHY?

- RIG AEROMECH TESTS GEARBOX COMPON. REVEL. - DES. & FAB.

IS P APF ROACH SO - FABRICATION METHODS RESEARCH M ND" - RIG TESTING 5 Ti BORSIC SHAFT - DESIGN & FAB. RESEARCH -RIG TEST [4^, 6 —1 PACKAGED PROTOTYPE FUEL CONTROL- DES. & BENCH L" flMPE TITI -1ST ENGINE RUN 7 StiOULD, ENGINE DEMO ENGINE DESIGN INTEGRATION & PROCURE E3E slult.

-ASS'Y. & INSTRUM.; IST RUN LD - DEMO ENGINE IMPROVEMENT - DEL. TO NASA 7, _.__.

_... . .-_ -. - .. _ ._... _. _.. --..--.-^. '^_'•^°;`?. '^^^.

r 'ins TABLE I. TASK I AIRCRAFT TYPES 1976 DATA 1988 MISSION POWER RANGE GROSS WEIGHT RANGE CRUISE TAKEOFF CRUISE RANGE KW KW NO. TYPICAL PAX PRICE RANGE SALES KG (NMI) WAS) (FT) CAT. EKG. PRODUCT PAYLOAD (HP) (HP) (LBS) (1000`s) 2 926 61.7 2440 1 (a) 1 16-31 2387 150, PA 18 56-67 74.6-89.5 697-719 SPORT 19 (1600-1650) (120) (8000) (75-90) (100-120) (500) II 1 -24-66 7246 172, BEECH C23 4 100.7-11.1.9 134-149 1089-1198 1297 87.4 3048 SIERRA B24R (135-150) (180-200) (2500-2750) (700) {170) (10000) CESSNA 206 159.6-168 1 212.3-224 1568-1655 1575 97.7 3048 IIIU 1 a 46-91 217.1 4`-.6.. (285-300) (3600-3800) (850 (190' (10000) PA32R LANCE 214-225 n^ 210 TC A36 IIIP 4=b' --- --- --- 166 (900) (210) .(18000) 2224 136 5486 IV 2 90-330 1484 SENECA, BARON 6-8 303 - 380 425-462 2396-2675 (40C-510) (570-620) (5500-6140) (1200) (265) (18000) B60 2 200-1,400 1083 6-12 820-1044 820-1044 4138-4487 3335 134 6095 V (b) SHRIKE, 421 CITATION (1100-1400) (1100-14001(9500-10300) (1800) (270) (20000) VI AG 1 40-80 1111 THRUSH 907 KG 166-417.6 224-596.5 1496-2721 16.2 KS 56.6 0 CESSNA A1888 2000 LB (225-560) (300-800) (3300-6000) (4.5 HRS) (110) RELIC 1-3 100-900 1030 NEW PRODUCT 2-5 6115 1 56.6 0 _ --- - - --- (330) (110) OPPORTUNITY 3250$ (a) Eliminated By Cast, Turbine Power Increased Aircraft Price 15-30% (b) Eliminated - Power Requirement Beyond Scope of Study TABLE II. INITIAL TURBOPROP ENGINE PERFORMANCE POSTULATIONS AT CRUISE TURBINE INLET TEMPERATURE AT CRUISE(a) CRUISE CONDITIONS 1367 0 K (20000 F) 1422 O K (21000 F) _ CATEGORY SPEED ALTITUDE POWER PRESSURE FUEL FUEL M/S M REQUIRED RATIO CONSTUMPTION PRESSURE CONSUMPTION

(KTAS) (FT) KW m /J RATIO m g/J

(HP) LB/HR-HP

LB/HR=HP

56.6 2440 81.8 79.4

w

I 8'8 g'6 (110) (8000) (90) (0.485) (0.467)

w

2440 1119 82.0 79.9

74.6

TT

8'7 g'4 (145) (8000) (156) (0.486) (0.473)

103 6096 167.8 74.4 72.2

III IO.O

(225) (200) (20000) {0.440) IO'8 (0.427)

134 6096 179.0 73.4 71.3

IV

9'8 10.6 (260) (20000) (24O) (0.434) (0.422)

157 6096 559.5 MO 70.8

V

IO'4 (305) (20000) (0.429) (0.491)

g'6 (750)

32509

(a) Maximum Temperature rating - 15330 K (23000 F)

TAS E 111. INITIAL TURBOPROP ENGINE POI-)ER.AND.AIRFLOW AT SEA LEVEL STATIC TURBINE 'INLET TEMPERATURE RATING AT CRUISE REQUIRED POWER 186.7 0 K (2000 0 F) 1422+ K (2100 0 F) (a) -SEA lEVEL STATIC PARAMETERS CATEGQRY CRUISE TAKEOFF AVAILABLE AVAILABLE POWER •AIRFLOW POWER AIRFLOW

K14 KG/S KW KC /S

(- H p ) (LB/•SEC) (HP) (LB /SEC)

67.1 89.5 116.3 0.34 100.7 0.29 I (90) (120) (156) (135) (0.74) (0,64,) I. I 111,9 149.2 194.0 0.56 167,1 0.48 (1.50) (200) (260) (1.23) (22 4) (1,08) 67.8 223..8 350.6 1.01 305.9 0.88 X 225) (300) ( 47.0) '(2.22) (410) (1.94) .

IV

I79.0 223, 3

.361,3. 1..04 314.8 0,91

(240) (300) (484) (2.29) (422) (2.00) V 55.9.5 559.5 3081.7 3:,3:2 947.4 2.73 (750) (1450) (6,88) {1270) (6.02) (750) -

32510

() MAXIMUM TEMPERATURE RATING - 3533 ) K (2300 c 0 F) TABLE IV, INITIAL TURBOFAN ENGINE PERFORMANCE POSTULATIONS BYPASS R 1 6:1 CRUISE CONDITIONS SEA LEVEL STATIC CATEGORY REQUIRED FUEL AVAILABLE FUEL AIRFLOW SPEED ALTITUDE THRUST CONSUMPTION THRUST CONSUMPTION KG/s mg /NS KN ing/NS (LB/SEC) M/S M KN (LB/HR-LB) (LB) (LB/HR-LB) (KTAS) (FT) (LB) 1.62 11.9 6.4 56.6 2440 1.04 14.5 (363) #.; I . (110) (8000) (234) (0.512) (0.42) (14.0) 2440 1.31 15.0 2.12 11.9 8.0 74.6 (145) (8000) (295) (0.530) (4.76) (0.42) (17.7) 1.39 15.7 3.45 11.9 12.5 103 6096 It (313) (0.553) (776) (200) (20000) (0.42) (27.5) 134 6096 1,16 17.2 3,02 11.9 10.9 4 (260) (20000) (261) (0.607) (679) (0.42) (24.0) 157 6096 3.11 18.3 8.41 11.9 30.2 11 llr ' (20000) 1 (700) (0.647) (1890) (0.42) (66.s (305) ,^ 32511 ORIGIAIAL ^k(a y p^® PP F iia wit 6k 61" 1 4 iotJ ^rtiYidu rind t4u^ TABLE V. BASELINE TURBOPROP SCALING DATA POWER (a) DIMENSIONS MM (IN) WEIGH:' KW KG (HP) A 8 C D E G (LBS) F H J K 713.5 361 110 230.4 346.2 440.2 743.7 176.8 322.6 J = H K =^ 80.4 (484) (4.33) (7.89) (13.63) (17.33) (28.09) (29.28) (6.96) (12.7) (177.0) ti EXP EXP EXP O. EXP EXP EXP EXP EXP EXP Ln _ `^ 0.12 0.27 0.27 0,27 0.27 0.50 0.36 --^ --- 0,72 SCALING RANGE 186 KW 119 KW TO (250 1 HP). (1500 HP) 116 91.4 157.5 279.4 330.2 561.3 584.2 100.3 254.0 49.0 (156) (3.6) (6.2) (11.0) (13.0) (22.1) (23.0) (3.95) (10.0) (108) 32512 (a) Power Available: Sea level Static DIMENSIONS MM (IN) THRUST) WEIGHT KN KG (LB) A B D C E P G H (LB) 216.4 .3.45 446.8 327.7 361.7 493.5 129.8 586.7 620.5 87.9 (776) (17.59) (12.90) (14.24) (19.43) (5.11) (8.52) (23.1) (24.43) (193.6) --- EXP EXP EXP EXP EXP! EXP EXP EXP EXP 0.50 0.50 0.36 0.42 0.40 0.40 0.33 0.33 0.80 COMPONENT RELATIVE COST: BASELINE TOTAL GEARBOX COLD HOT ACCY & COMPRESSOR COMBUSTOR TURBINE A & T HSG HSG SYSTEM AIR INLET 0.20 0.16 0.06 0.20 0.02 0.08 0.24 0.08 1.0 GATE: ADVANCED TECHNOLOGY 0.20 0.08 1 0.04 1 0.07 1 0.01 0.04 1 0.12 1 0.04 1 0.-6 FABRICATION TECHNOLOGY 0.048 1 1 0.16 1 0.04 1 0.042 1 0.01 0.04 0.12 0.04 0.5 TABLE VIII. GENERAL AVIATION MARKET AND GROWTH TREND SUMMARY 1988 MARKET ACTUAL 1976 SIMPLE DELIVERIES UNDISTURBED DISTURBED ANNUAL GROWTH CATEGORY RATE - % PISTON TURBOPROP PISTON TURBOPROP UNDISTURBED PISTON TURBOPROP f 2,387 I 0 3,100 0 3100 0 2.5 II 7,246 0 12,000 0 2400 9,600 5.5 III 2,171 0 3,210 640 2,570 4.0 I-III 11,804 0 18,310 6140 4.6 0 12,170 1,484 0 IV 2,230 0 1100 4,400 4.2 638 810 0 3,000 V 473 990 5.2 I-V 13,926 473 21,350 990 7240 19,570 4.6 TOTAL 14,399 22,340 26,810 AG 980 . 1 0 1,500 1,500 4.4

32515

TABLE IX. PROJECTED LIGHT HELICOPTER MARKET WITH GATE ENGINES FOR THE 5 YEAR TIME FRAME FROM 1988 TO 1993 NUMBER NUMBER HELICOPTER DESIGNATION HELICOPTERS OEM ENGINES SINGLE 1700 1700 TWIN 700 1400 TRI-PAC 350 1050 TOTAL 2750 4150

32516

TABLE X. MISSIONS FOR GATE POWERED LIGHT HELICOPTERS u SINGLE TWIN TRI-PAC s AGRICULTURAL ® CORPORATE CORPORATE EXECUTIVE EXECUTIVE * SEARCH AND a SEARCH AND ® SHUTTLE RESCUE SERVICE RESCUE ® POLICE ®POLICE &AMBULANCE o TRAINING AMBULANCE *OFFSHORE ® PHOTOGRAPHY a OFFSHORE O AGRICULTURAL SINGLE TWIN TRI-PAC 261 KW 522 Kai 783 KW

(350 , IP) (.700 HP) (1050: HP)

GROSS KG 1271 1 952 3337 WEIGHT (LBS) (2800) (4300) (7350) EMPTY;. KG 726 976 1680

WEIGHT (LBS) (1600) ('2150) (3700)

f

UttFUL KG 976 1657

LOAD ` (LBS.) (1200). ('.2150) (3650): 3: 8 NUMBER SEATS 5 182 431 70 KG 4 FUEL (LBS) (400). (950) (1550)

t' g/U 101 93 93

SFC (LB/HP-HR) (0.60) (0.55) (0.55) @POWER RATING 75% 60% 60% 33 834 1019 RANGE KM 4

(rim) (180) (450) (550)

` RESERVE- MIN 30 45 CRUISE M/S 46.3 72.0 77.2 SPEED (KNOTS) (90) (140) (150) (MPH) (104) (161) (173) SERVICE m 3658 (12 000/5000S a) (147000/8,000/4,000) CEILING FT (12000).

HOVER, OUT OF M 1829 2438 1219 GROUND. EFFECT FT (6000) (8000) (4000) ^.

CABIN 0.10 0.05/0•.08 0.05/0.08 VIBRATION C ABIN 90 75/80 75/80 NOISE dB2 SELLING 100,000 to 300,000 to 700,000 to $ t PRICE 125,000 500,000 12000,000

TABLE X1I. 1988 GATE POWERPLANT CAPABILITIES AND REQUIREMENTS

SALES/ REL. REL. COST

REL. INSTALLED MULTI AVG. SERVICE

POWER/ DEVELOP. PRODUC

ENGINE TYPE POWER/WT. CRUISE FUEL IN

TEO FRONT AREA CAPITAL TION SFC ?

PLACE?

RECIP. - SPARK 1003 100 100 (a) LOWEST LOW NO <2000 YES ('I 976=110) GAS TURBINE 200.320 280-450 110-130 MED. MED. YES X4000 YES DIESEL (POTENTIAL) 100 MED. ?

100 80-90 (b) HIGHEST YES PART ROTARY 180-220 200-300 HIGH ? ?

96-120 YES (b) NO

32519

(a) COOLING CAPABILITIES UNKNOWN (b) CURRENT EXPERIENCE IS LIMITED

TABLE XIII. GATE EMISSIONS, EPA 1979 STANDARDS LTO CYCLE'

ENG. CLASS CO (a) THC (a) NOx (a) SMOKE (Ii)

PROD. STD. PROD. ISTD. PROD. STD. PLOD. STD.

P1-PISTON (c) 50-120 42 3.0-4.5 1.9 0.2-1.3 1.5

132-TURBOPROP 20-30 26.8 6-12 4.9 6-10 12.9 -._ <50

T1-THRUST

15-60 9.4 4-16 1.6 3.7 _ <32

2.5-4.5

<3629 KG (8000 LS)

0.9-1.2 3.3-3.4 INVISIBLE

OCGAT (013J.} 6.9-7.2

32520A (a) KG/1645 KW-HRS/CYCLE (LB/1000 HP-HRS/CYCLE) FOR PIST014 ENGINES AND TURBOPROPS, KG/9.81 KN THRUST-HRS/CYCLE (LB/1000 LBS THRUST-HRS/CYCLE) FOR THRUST ENGINES (b) RELATIVE REFLECTIVITY (REFERENCE 15) (c) P1 CLASS E)(CLUDES RADIALS (REFERENCE 15)

t

S F TURBINES RECIP'S POWER-SLS AND TAKE-OFF TURBINES OEM ENGINE CATEGORY TURBINES SALES —1 COMPANY TOTAL GATE (a) KW AIRCRAFT Et (HP) SALES (50% OF MARKET) ENGINE SALES II 175/198 1-2000 4800 (235/265) 205/283 IiIU (2731380) 3210 12$5 2570 238/421 111P (3201565) 220/395 1V (2951530) VI 1500 750 (;400/480) (AGRICULTURAL) h 261 ± 56 HELICOPTER 830 (350 t 75) 550 SAME 23300 TOTAL 12055 4220 8150 SPARES 16285 31450 GRAND TOTAL $120 (b)' $220 (b) MARKET VALUE, /a rtoo not i n no% II& IIIU III P'& IV CATEGORY HELICOPTER 92.6 123.5 56.6 M/S.

SPEED (KTAS). (110) (180) (240) 3048 5486 m ALTITUDE ( (FT). SEA LEVEL (10000) (18000) R COMPRESSO PRESSURE RATIO 9 14 20 EFFICIENCY - PERCENT 82 82 82 PARAMETRIC RANGE OF EFFICIENCIES EVALUATED 82/80 79.5/77 , 75/72 FOR SENSITIVITY-PERCENT COMBUSTOR EFFICIENCY PRESSURE LOSS FUEL HEATING VALUE 3.5 PERCENT 42 798 U/g (18400 B/LB) 99.5 PERCENT TURBINE: T CRUISE = T MAX - 167 0 K (300 0 F) T CRUISE ROTOR NOZZLE, ROTOR $ SHROUD °K TYPE EFFICIENCY AND SEAL COOLING ( 0 F) PERCENT PERCENT 1256 COOLED 85 3.0 (1800) UNCOOLED 87 1.0 1339 COOLED 84 5.0 (1950) UNCOOLED 87 2.0 8.0 1422 COOLED 82 4.0 (2100) UNCOOLED 87 SENSITIVITY RANGE: EFFICIENCIES VARIED ±2% CRUISE TURBINE (a) ASSUMPTION EQUIVALENT BYPASS COOLING INLET TEMPERATURE BLEED AIRFLOW (NOZZLE, ROTOR AND SHROUDS)

d

PERCENT (e^ 1.0 1256 OPTIMISTIC - CONSER 3.0

(1000) VA TIVE

d " 1339 OPTIMISTIC<. 2.0 (1950) CONSERVATIVE 5.0 1422 OPTIMISTIC 4.0 (2100) CONSERVATIVE 8.0 32525 (a) Maximum Rated Temperature is, 167 0 K (3000F) Greater Than Cruise TABLE XVII. DATA MATRIX FOR THE TURBOFAN PARAMETRIC ANALYSIS

a

BYPASS RATIO 5 TO 11.1 s.

FAN P.R. 1.20 TO 1.60 @ 89.5% EFFICIENCY CORE P.R.

7 TO 13.1 @ 83 TO 80% EFFICIENCY CRUISE T.I.T.

1037 TO 1365 °K (1400 TO 2000°F) CORE TURBINE EFFICIENCY .

88% FAN TURBINE EFFICIENCY 89% SERVICES : 1% PRESSURIZATION BLEED PLUS 2.24 KW (3 HP) POWER EXTRACTION OTHER CYCLE VARIABLES = CONSISTENT WITH 1985 TECHNOLOGY -- COMBUSTOR EFFICIENCY = 99.5% @ 3.5 % PRESS.

LOSS CORE TURBINE EFFICIENCY = 88% FAN TURBINE EFFICIENCY = 89%n , SHAf7 MECH.. EFFICIENCY = 99% (BOTH) PRIMARY DUCT P = 3.0% DUCT MAC4NO. = 0:33: FAN DUCT P = 4.0% (SEE TEXT) NOZZLES Cf = 98.5% 32420 112.

TABLE XVIII. TYPICAL PERFORMANCE AT 107 KN (240 LB) THRUST ^- OPTIMIUM SFC LINE SPEED CORRECTED TURBINE SPECIFIC SPEED INLET FUEL CONSUMPTION TEMPERATURE O K LB % OF MAX % OF DESIGN ( O R) mac NS HR-LB 100 104.9 1265 (2280) 13.17 (0.465) 95 99.7 1226 (2210) 12.52 (0.442) 90 94.7 1376 (2480) 13.34 (0.471) 32526 f TABLE XIX. 1985 STATE-OF-THE-ART COMPRESSOR SUMMARY AT A CONSTANT AIRFLOW OF 1.19 KG/S (2.62 LB/SEC) FLOWPATH ADIABATIC SPEED CENTRIFUGAL PRESSURE LENGTH RADIUS NO. CONFIGURATION TIP TIP RATIO EFFICIENCY REAM) SPEED WIDTH MM lim (RPM) PERCENT M/S MM (IN) (IN) (FT/SEC) (IN) 1160.3 640.1 5.72 82.6 165.1 C 81.5 (69 900) (2100) (0.225) (3.25) (65b) 780.9 487.4/457.2 8.79/5.33 344.8 190.5 C "- 80.4 9:1 (47 040) (1599/1500) (0.346/0.210) (5.70) (7.50) 1143.7 587.0 4.72 129.5 154.9 --_ AC 81.8 .(68 900) (1926) . (0.1$6) (5AO) (`6;:10) 90.. 4 1143.7.. 472.7 6.65. 205.7 527.0 (68 900) (1551) (0.262.) {5,00).:, (8.10) 1367.0 624.2 5.25 154.9 14262 11.3:1 3010 AC 81.7 (82 350) (2648) (0.206) (6.10) (5.60) 1367.0 670.6 5.23 165.1 147.3 14:1 -- AC 79.2 (82 350) (2200) (0.206) 16.50) .(5.80) 1367.0 563.9 3.99 208.3 134.6_ 15:1 4010 AAAC 78.2 (82 350) (3850) (0.157) (8.20) 1367.0 640.1 2.84 2083 137,2 --- AAAC 74.3 (82 350) (2100) (0.112) (8.20) (5.40) 20'1 1246.3 670.6/487.7 5.92/2.51 177.8 190.5 5DI0 CC 76.6 (75 082) (2200/1600) (0.233/0.099) (7.Op) (7."50) CONFIGURATION C. AC11.3 AAAC 15 CC 20 COMPRESSOR 9.0 11.3 15.0 19.8 PRESSURE RATIO COMPRESSOR EFFICIENCY } 81.5 81.7 78.2 76.6 PERCENT 1327 1254 1257 TURBINE INLET °K 1249 TEMPERATURE (1930) (1800) (1790) (1805) ( O F) BYPASS COOLING 2.0 1.0 1.0 1.0 PERCENT TURBINE EFFICIENCY 87.9 88.6 88.0 88.0 PERCENT REVS 1167 1375 1375 1381 SPEED (RPM) (69900) (82500) (82500) (82900) CORRECTED KG/s 1.19 1.19 •1.19 1.x.9 AIRFLOW (LB/SEC) (2.62) (2.62) (2.62) (2.62) COMBUSTOR PRRSSURE = 3.5% JET NOZZLE = 1.17 DROP PRESSURE RATIO FUEL. LOWER = 42 798 J/ PROPELLER EFFICIENCY = 86% HEATING VALUE (18 400 B/LB^ EXHAUST DUCT — 3% PRESSURE LOSS COMBUSTOR EFFICIENCY = 99.5% MECHANICAL EFFICIENCY= 98.5% 32530A TABLE XXI. CANDIDATE ENGINE CRUISE PERFORMANCE: CORRECTED AIRFLOW AT SEA LEVEL STATIC, 1.19 KG/S (2,62 LB/SEC) CATEGORY IV CATEGORY III P CATEGORY III U CATEGORY II 128.6 M/5 250 KTAS) 108.0 M/S (210 KTAS) 97.7 M/S (190 KTAS) 87.4 N/S 170 KTAS @ 5486 M 18000 FT) @ 5486 M (18000 FT) @ 3048 M (10000 FT) @ 3048 M 10000 FT POWER FUEL POWER POWER POWER FUEL FUEL FUEL AIRFLOW CONSTUMPTION AIRFLOW CONSUMPTION AIRFLOW CONSUMPTION AIRFLOW CONSUMPTION CONFIGURATION POKER KW POWER KW POWER KW POWER KW uc! _ lug tj _ μ 9 KW 1 J 96 7 - 5 KW KG S J KW R KW J (HP) ( HP) HP N ( (( ( ( 1 j 1 j^ f 1 E^ ( HP) b/SECT I HRL EHPI LLB / SEC l R-EHP1 \L8 H5EC / \ Rif -EHP/ 1^^] (HP) -»WHO l

N

rn 76,8 C9 200 286 191 281 '' 78.5 233 260 81.7 228 25B 82.8 (26B) (174) (0,455) (256) (171) (0.465) (312) (158) (305) (157) (0.40) (0.484) AC 11.3 183 271 71.9 176 260 73.5 213 238 76.8 208 235 77.5 (246) (163) (0.426) (236) (158) (0.435) (285) (145) (0.455) (279) (143) (0.459) 163 234 AAAC 15 73.1 157 232 74.3 184 205 79.2 180 204 80.0 (219) (142) (0.433) (211) (141) (0,440) (246) (125) (0.469) (242) (124) (0.474) CC 20 149 212 73.6 144 212 74,8 163 182 81.2 160 182 81.9 (200) (129) (0.436) (193) (129) (0.443) (219) (111) (0.48.L) 1 (215) (111) (0.485) 32531A POWER FUEL AIRFLOW TEMPERATURE POWER CONSUMPTION CONFIGURATION KW 0 KV OF) (EHP) EHP (LB/NR-^EHP) T' - B/ 1504 362.4 304.1 86.0 C 9 (2250) (486) (185) (0.509) 1432 344.5 289.3 79.4 AC 11.3 (2120) (462) (176) (0.470) 1426 312.4 261.4 79.8 AAAC 15 (2110) (419) (159)- (0.472) 1435 290.1 243.3 79.4 CC 20 (2125) (389) (148) (0.470) DIMENSION - rrm (in) WEIGffT d GATE KG y A C(a) D E F G K DESIGN B (LS).

EQUAL;PDWER ::365.5 XW (490 HP)

713.5 340.4. 457.2 147.3 93.5

310 200.7 ---- 440.2 850.9 182.9

(4.33) (7.89) (17.33) (28.09) (33.5) (18:43) (:S).. (206)_:

(7.2) (13:4)

188 309:9, 99:'4

840.7 942.2.

(219)

(33.1) (37.1) (7.4) (12.2)

894.1 995. 7 1.98,.1 322.6

.4010

-(240)

(35.2) (3,' 9. 2) (7.8) (12.7)

424.2 1 2

811.2 972.8 205,7

:(8.3) (16.7) (268,) .

(34,3)__ . (38.3)

co

/SdC) EQUAL ANFLOW: 3. ? 9 KGJS4:C (246Z LB

713.5

850.9 3.82.9 340.4 93.5

(33.5) (206)

(28.09) (13.4)

833.1 97.6

934.7 182.5 304.8

(215).:

(32.8) (36.8) (7.2). (12.0)

4010 871.2 972.8 182.9 304.8 302.6.

(226)

(34.3) (38:3) (7.2) (12.0.). ..

i

838.2 939.8 1829 391.2 109:9

(33.0) (37.0) (15.4) (242):..

(7. 2)

i SCALING EXPORENT

4.36 04.6

0.32 1 0.27 0.27 0.27 0.27 0.50 0.36 0.72

i

(a) NOT USED

SHAFT SUPPORT STIFFNESS CRITICAL SPEEDS FRONT FIRST SECOND THIRD REAR

MN/m rev/.s -rev /s

rev/s; MN/m

(LB/`iN) (LB/IN) (RPM) (RPM (RPM)

175 217 1530 2770

(106) (x.06)

(13024) (91804) (166230)

32534 TABLE XXV. RELATIVE COST SUMMARY FOR FOUR BASIC ENGINE CONFIGURATIONS COMPONENT RELATIVE COST: BASELINE TOTAL GEARBOX COLD HOT ACCY COMPRESSOR COMBUSTOR TURBINE A & T HSG. HSG SYSTEM AIR INLET 0.20 0.16 0.06 0.20 0.02 0.06 0.24 0.06 =1.0 GATE 2010 0.20 1 0.08 0.04 0.07 1 0.01 0.04 1 0.12 0.04 0.6 1 , r; GATE 2010 WITH FABRICATION TECHNOLOGY 0,16 0.048 0.04 0.042 0.01 0.04 0.12 0.04 0.5 r t.

GATE 4010 0.20 0.32 0.04 0.20 1 0.04 10.06 0.12 0.08 1.06 GATE 4010 WITH FARBICATION TECHNOLOGY 0.16 1 0.192 1 0.040.12 0.04 0.12 0.06 0.08 0.812 IT GATE 6010 0,20 0.16 0.04 0.20 0.02 rO.O6 0.12 1 0.06 1 0.86 GATE 5010 WITH FABRICATION TECHNOLOGY

0.16 1 0.096 1 0.04 1 0.12

1 0.02, 0.06 0.12 a.06

0.676

TABLE XXVI. OEM COST FOR THE FOUR BASIC ENGINE DESIGNS: } EXPRESSED IN DOLLARS FOR A PRODUCTION RATE OF 500 PER YEAR f CONSTANT WEIGHT CONSTANT POWER DESIGN NUMBER 227 KG (500 LB) 365.5 KW (490 HP) 2010 59870 33115 3010 77831 44588 4010 105770 63970 5010 85814 55552 FAN DIAMETER ................. ............. 311.2 mm (12.25 IN) BYPASS DUCT DIAMETER .... . .................... 386.1 mm (15.2) FAN PRESSURE RATIO ......................................... 1.25 FAN SPEED .......................... . . 311.7 rev/s (18700 RPM) POWER ......................................... 197.7 KW (265 HP) THRUST .......................................... 1.73 KN (388 1-8) COMPONENT RELATIVE COST. BASELINE TOTAL .GEARBOX COLD HOT AGCY & COMPRESSOR COMBUSTOR TURBINE A & T HSG HSG SYSTEM AIR INLET 0,20 0.16 1.0 0.06 0.20 0.02 0.06 0.24 0.06 GATE 2011

0.34 0.08 0,04 0.07 0.02 1 0.04 0.12 j 0.04 F 0.75

GATE 2011 WITH FABRICATION TECHNOLOGY 0.27 1 0.048 0,042 1 0.02 1 0.04 1 0.12 1 0.04 1 0.62 1 0.04 1 GATE 2012• 0.30 1 0.08 0.04 0.07 1 0.02 1 0.04 1 0.12 0.04 0,71 GATE 2012 WITH FABRICATION TECHNOLOGY 0.24 0.048 1 0.04 1 0.0421 _ 0.02. 1 0.04 1 0.12 1 0.04 1 0.59 32538 TABLE XXIX. COMPARISON OF A DIFFERENTIAL TURBINE DESIGN TO A SINGLE SHAFT DESIGN SINGLE SHAFT DIFFERENTIAL TURBINE DESIGN NO. 2010 2011 2012 y APPLICATION TURBOPROP TURBOPROP TURBOSHAFT OUTPUT FLANGED FLANGED SPLINTED SHAFT 33 rev/s (2000 RPM) 33 rev/s (2000 RPM) 100 rev/s (6000 RPM) COST % 100 144 131 WEIGHT % 100 126 119 SFC % 100 104 104 32539 121.

^ 111 U III CATEGORY I °I P IV K0 363 454 464 544 PAYLOAD (LB) (800) (1000) (1200) (1000) CRUISE M/S 10 8.,0 128.6 87.4 97.7 SPEED (KTAS-) (210) (250) (170) (-190) CRUISE 'M 3048 3048 -54486 5486 ALTITUDE (FT) (10000) ' "(18000) (18000) (10000) KM 1296 .1574 1667. 2222 RANGE (NM) (700) (850). (900) .(12.00) T. 0. DISTANCE P S I 488 ;610 610 671 SL.: ISA (2200) (FT) (,1600)- (2000) (2000) LANDING DISTANCE M 427 457 549 SL: TSA (FT) (1400) ;05.00) ,(1500 ) (1800) ASPECT RATIO 8 8 9 1 9 CL TAKEOFF 1.:6 1..6 1.6 1.8 MAX 2.3 C LANDING 2.1 2.2 2.2 LMAX 85 85 PROPELLER EFFICIENCY 85 85 @ CRUISE -- PERCENT CRUISE POWER (a) LAPSE RATE SLTO 0.641 0.656 0.508 0.535 79.6 CRUISE LB Kge3 r_ 82.4 81..2 81..7 SFC \HR-EHP (0.488) (0.481) (0,484) (0.471) RESERVE FUEL (b) S 60 60 60 (1) (1) @CRUISE POWER (HR) (1) (1) 32540 (a) SLTO: Sea Level Takeoff Thermodynamic p ower. Gearbox Torque Limited to Lower Rating (b) Include Takeoff Reserves and Climb Allocations,per Beech Experience in Similar Designs i TABLE XXXI. BASELINE AIRCRAFT POINT DESIGN CHARACTERISTICS CATEGORY III U III P IV II TAKEOFF 1313 1607 1837 3127 KG WEIGHT (3543) (4049) (6894) (LB) (2894) FUEL 192 298 333 703 KG WEIGHT (LB) '(424) (658) (734) (1549)

WING M2 13.7 15.9 18.4 20.2

AREA (FT2) (148) (171) (198) (217) 13.5 WING M 10.5 11.3 12.9 SPAN (44.2) (FT) (34.4) (37.0) (42.2) CL CRUISE 0.23 0.26 0.23 0.25 KG USEFUL .752 787 1247 '^' (a) LOAD (LB) (1224) (1658) (1734) (2749) SLTO (b) KW 198 284 421 395 (530) (HP) (265) (381) (565) CRUISE POWER KW 189 220 128 215 REQUIRED (d) (HP) (172) (253) (295) (288)(c) ACCESSORY KW 1.5 ).0 3.0 2.2 POWER (HP) (4)(c) (2) (3) (8) C TAKEOFF POWER 174 202 234 218 KW REQUIRED (234) (271) (314) (292)(c) (HP) EMPTY 758 855 1050 1880 KG WEIGHT (LB) (1670) (1885) (2315) (4145) PROPULSION (e) KG 136 154 187 419 WEIGHT (LB) (299) (339) (412) (923) [ INDUSTRY COST PER AN UAL KG (LB) OF AIRFRAME SALES GROSS RETAIL CATEGORY QUANTITY PRICE WEIGHT (UNITS) () () II 9,600 46.30 (21) 52,700 IIIU 21000 48.50 (22) 65.,400 III-P 2,000 90-.40 (41) 131,300 IV 3,500 90.40 (41) 209,500 32542 (a) All Price Estimates are in 1976 Dollars to be

Consistent with the Terms of the Task I Market

Survey TABLE XXXIII. AIRCRAFT PARAMETERICS - LINEAR INFLUENCE COEFFICIENTS a a a PAYLOAD a SFC FUEL CRUISE CRUISE FUEL POWER POWER TOGW WEIGHT TOGW WEIGHT KW/KG KC/%ASFC KW/%ASFC CATEGORY KG/%ASFC (EHP/LB) (LB/%ASFC) (EHP/%ASFC) (LB/%QSFC) 5.87 2.14 0.197 0.556 II 3.43 0.315 (I:2.94) (4.73) (0.120) (0.746) 0.227 5.57 3.08 0.543 III U 3.27 0.338 (0.138) (6.8) (0.728) (18.9) 0.247 11.73 4.97 0.483 3.363 III P 0.350 (0.150) (25.87) (10.95) (0.647) 0.222 15.92 36.38 1.907 IV 4.05 0.718 (0. 135) (80.2) (35.09) (2.556) 32543 TABLE XXXIV. AIRCRAFT SYNTHESIS ANALYSIS RESULTS, VARIATIONS IN CONFIGURATION AND RETAIL PRICE DUE TO ENGINE CHANGES (CONSTANT MISSION).

CATEGORY III P IV (e) II BASELINE ADJUSTED BASELINE ADJUSTED BASELINE ADJUSTED ENGINE CONFIGURATION C9 (b) iv tc) AC 11. 3 C9 (b) Co (c) AC 11.3 C9 (b) C9 (c) AC 112 1296 1837 1746 TAKE OFF KG 1313 1315 1789 3127 3008 2803 {4049 (3849) (6894) (6834) GROSS WEIGHT (L85) (2894) (2899) (2858) ► {3944) (5180) EMPTY KG 758 759 752 1050 1022 1007 1880 1817 1704 WEIGHT (Les) (1670) (1673) (1658) {2315) (2253) (2220) (4145) (4007) (3757) S63 454 454 454 544 544 544 KG 353 383 PAYLOAD (1000) (1000) (1000) (1200) (1200) {1200) (LBS) (600) (600) (800) 647 555 FUEL KG 192 193 181 333 313 285 703 WEIGHT (LBS) (424) {428) (400) (7343 (691) (529) (1549) (1427) (1223) 60 60 65 106 105 99(d) 10gd) ENGINE KG 10,1d) 113 WEIGHT (LBS) (133) (133) (144) (233) (232) (250) (221) (2181 {234) 184 419 416 430 KG PROPULSION KG 136 136 141 187 ISO (e (299) (310) (4121 (411) (429) (923) (917) (959) WEIGHT (299) CRUISE KW 128 128 126 220 218 215 21§d) 2111 20jd) 1283)1d POWER (EHP) (172) (172) (169) (285) {293) (2881 (288 ► (277) CRUISE LB 82.4 82.6 7T.7 81.7 78.5 735 7816 76.8 72.0 Me^J (0.460) (0.484) (0.485) {0.435) (0.471) (0.455) (0.426) SFC HR-EHP (0.488) (0.490) AIRCRAFT (1) 1977 139,000 233,300 222,500 210,600 62,500 62,500 64.600 143,800 139,300 RETAILPRICE DOLLARS 1 11 (a). Min engine aircraft, others are single engine 33883 ( p ) hosed on task I engine data (c) based on revised task 11 engine data (d) each engine ( ( er(s) and starter generators) (e) propulsion weight includes engine ( s), controls, exhaust pipe(s), all system with cooler s), fuel system, propel (f) Includes engine price TABLE XXXV. L C 2 FINAL SCHEDULE DEFINITIONS CA T. II IIIU IIIP IV INSURANCE/YR. $ 2120 2625 3126 4240 HANGAR/YR. $ 900 1000 1600 2600 AIRCRAFT + PROP.

4.10 7.61 7.61 SERVICE/HR. $ 4.10 ENGINE SERVICE/HR.

MTCE BASIC W $8/HR.

TBO @ 3500 HRS.: 30% OF ENG. RETAIL PRICE HOT SECT. INSP.: IGNORE (INCL. IN MTCE.)

MISC. COSTS: FIXED/YR. 1% of Acq. -- 1.80 2.50 3,00 : VARIABLE/HR. 1.50 20% DOWN + 80% FINANCED @ 1.0%, 5 YRS, ACQUISTION: 30% RESALE = 0.92 x RETAIL PRICE

32545

TABLE XXXVI. COMPARISON OF GAVE 1988 CATEGORY 11 AIRCRAFT WITH TURBOPROP -vs- RECIPROCATING ROWER.: CONSTANT AIRCRAFT SIZE; VARIABLE PERFORMANCE CATEGORY II WITH RECIPROCATING CHANGE ENGINE GATE RELATIVE TO CONSTANT TO WEIGHT & CRUISE CATEGORY Ii GATE ENGINE POWER DELTA PERCENT 174.4 174.4 ENGINE TAKE-OFF POWER KW(HP) (234) -'-- --- (234) 1312.7 1312.7 MAX. T.A. WEIGHT KG(LB) (2894) (2894) 757.5 99..8 857.3 STANDARD EMPTY WEIGHT KG(LB) (1670). (220) +13 (1890) 455.4 555.2 99.8 -18 (1004) USEFUL LOAD KG(LB) (1224) (220) 142.4 192.3 49.9 (110) -26 (314) USABLE FUEL KG(LB) (424) 363 49.9 313 i PAYLOAD WITH FULL FUEL KG(LB) (800) (110) -14 _ (690) L,1 87.4 2.57 84.9 MAX. CRUISE SPEED m/s (kts) (170) (5) -3 (165) i^ (10000) ALTITUDE m(FT) (10000) ---- --- 12.6 126.8 (170) --- (170) CRUISE POWER KW(HP) J 1297 447 (700) (240) -34 (460) RANGE KM(NM) 487.7 TAKE-OFF DISTANCE OVER 487.7 (1600) ---- --_ (1600) 15.24 M (50 FT) 426.7 426,7 LANDING DISTANCE OVER 15.24 (1400) - -_ --- (1400) N (50 FT) - 13.75 ___. ___ (FT2 } (148) (148) WING AREA M2 i t TABLE XXXVII. COMPARISON OF GATE 1988 CATEGORY IV AIRCRAFT WITH TURBOPROP -vs- RECIPROCATING POWER: CONSTANT AIRCRAFT SIZE; VARIABLE PERFORMANCE CATEGORY IV WITH RECIPROCATING ENGINE CHANGE RELATIVE TO CONSTANT TO GATE WEIGHT & CRUISE CATEGORY IV GATE ENGINE POWER DELTA PERCENT 217.7 ENGINE TAKE—OFF POWER KW(HP) (292) --_— -- (^) i l 3127.1 3127.1 _ - (6894) ^ (6894) MAX. T.O. WEIGHT KG(LB) 2115.1 1880.2 234.9 (4145) (518) +12 (4663.)

STANDARD EM PTY WEIGHT KG(LB) 1247 234.9 1011.9 f i (2749) (518) -19 (2231) Q2 USEFUL LOAD KG(LB) 117.5 585.1 702.6 (1549) (259) -17 (1290) USABLE FUEL KB(LB) 426.8 544.3 117.5 (1200) (259) 1 -22. (941) PAYLOAD WITH FULL FUEL KG(LB) 8.2 120.4 128.6 r p i; (250) (16) -6 (234) MAX. CRUISE SPEED m/s (kts) .r 5486.4 5486.4 (18000) (18000) -`-_ --- ALTITUDE m(FT)

11.8

211.8 ~ (284) CRUISE POWER KW(HP) (284) ---- -- 560 1656 -26 (894) RANGE KM(NM) (1200) (306) 670.6 TAKE-OFF DISTANCE OVER 670.6 --_- _-- (2200) 15.24 M (50 FT) (2200) 548.6 LANDING DISTANCE OVER 548.6 L (1800) 15.24 M (50 FT) (1800) -`-- -- 20.2 _.-- (217) DING AREA M 2 (FT2 ) (217) ^ u- CATEGORY 11 WITH RECIPROCATING CHANGE ENGINE GATE RELATIVE TO CONSTANT GATE ENGINE MISSION CATEGORY II PERCENT DELTA 174.4 174.4 (234) ---- -^ (234) ENGINE TAKE-OFF POWER KW(HP) 1312.7 161.5 1474.2.

MAX: T. 0. WEIGHT KG(LB) (2894) (365) +12 (3250) 138.8 896.3 757.5 STANDARD EMPTY WEIGHT KG(LB) (1670) (306) +18 (1976) 555.2 555.2 USEFUL LOAD KG(LB) (1224) `-- (1224) 192.3 22.7 215 x-12 (474) USABLE FUEL KG(LB) (424) (50) PAYLOAD WITH FULL FUEL KG(LB) (800) --- (800) 87.4 87.4 •__ __ - MAX. CRUISE SPEED m/s (kts) (170) (170) 3048 3048 (10000) ALTITUDE m(FT) (10000) --^^ --- 126.8 14.9 141.7 (20) +12 (190) CRUISE POWER KW(HP) (170) 1297- (700) RANGE KM(NM) (700) M`-- --- TAKE-OFF DISTANCE OVER 487.7 73.2 560.8 15.24 M (50 FT). (1600) (240) +15 (1840) 426.7 LANDING DISTANT OVER 426.7 15.24 M (50 FT) ( 1 400) ---~ __ (1400) + ^ 13.75 1.21 14.96 WING AREA M 2 (FT . ). (148) (13) +9 (161) TABLE XXXIX. COMPARISON OF GATE 1988 CATEGORY Ili AIRCRAFT WITH TURBOPROP -v a s- RECIPROCATING POWER: EQUAL AIRCRAFT PERFORMANCE; VARIABLE SIZE CATEGORY IV WITH RECIPROCATING CHANGE ENGINE GATE RELATIVE TO CONSTANT CATEGORY IV GATE ENGINE MISSION DELTA PERCENT 217.7 99.2 316.9 (133) (425) ENGINE TAKE-OFF POWER KW (HP) (292) +46 3127.1 615.1 3742.2 (8250) MAX. T.O. WEIGHT KG (LB) (6894) (1356) +20 1880.2 474.0 2354.2 (4145) (1045) +25 (5190) STANDARD EMPTY WEIGHT KG(LB) 1247 141.1 USEFUL LOAD KG(LB) (2749) (311) +11 (3060) 702.6 141.1 843.7 USABLE FU"cL KG(LB) (1549) (311) +20 (1860) 544.3 544.3 PAYLOAD WITH FULL FUEL KG(LB) (1200) ---- "^ (1200) 128.6 128.6 MAX. CRUISE SPEED m/s(kts) (250) ---` --- (250) 5486 5486 ALTITUDE m(FT) (18000) ---- -- (18000) 211.8 41.8 253.5 (340) CRUISE POWER KW(HP) (284) (56) +20 RANGE KM(NM) (1200) "--` (1200) TAKE-OFF DISTANCE OVER 670.6 121.9 548.6 (400) (1800) 15.24 M (50 FT) (2200) -18 (a) 548.6 548.6 LANDING DISTANCE OVER -_- --- (1800).

15.24 M (50 FT) (1800) 3.9 20.2 24.1 WING AREA M2 (FT 2 ). (217) (42) +19 (159) 32549 (a) Performance is Equal Except for Take-Off Distance.

tr-.^ %GATE IMPROVEMENT CAT. IV CAT. 11 TOGW EMPTY WT. (PRICE) 18 CRUISE H.P. REQ.'.D 12 20 FUEL REQ'D —22 (ACCEPT.)

TAKEOFF DIST.

_ PAYLOAD, ALT., RANGE, VEL., EQUAL LANDING DIST.

02550 TABLE XLI. SUMMARY OF COMMON CORE APPROACH THERMODYNAMIC POWER(a) l..I GEARBOX CONFIGURATION RATING WEIGHT DESIGN.SHAVING RANGE COMMON CORE DESCRIPTION APPROACH —DESIG^SiSVOR (HP) {HP} (LB) (NP) e.

C9 BASIC 2010 DESIGN 365.5 93.5 365.5 ONE FRAME SIZE 198 — 422 PLUS 2010 PLUS WIDE RANGE (206) (490) (490) (265)—(565) SHAVING FLOWPATH SHAVING C9 BASIC 2010 DESIGN 365.5 93.5 365.5 317 — 422 TWO FRAME SIZE 2010 PLUS REDUCED RANGE (490} (206) (490) (425)—(565) FLOWPATH SHAVING PLUS :.1 C9 SCALED 2010 DESIGN 224 65.8 198 — 280 SHAVING 2010 PLUS REDUCED RANGE ^ 6 (300) (145) (300) (265)—(375) FLOWPATH SHAVING.

AC 11.3 3010 DESIGN 410 92.2 422 MODIFIED FOR TWO FRAME 3013 (550) (565)' ---- (203) COMPONENT REMOVP,L,.

FAMILY C9 205 2010 DESIGN 78.1 250 APPROACH DERIVED (275) (172) (335) ---- FROM 3013 DESIGN (a) SLS Turbine Inlet Temperature 15040K (22500F) 13 0 E k TABLE XLII. TWO FRAME FAMILY PERFORMANCE SUMMARY (SEA LEVEL STATIC, UNINSTALLED) s r FUEL AIRFLOW ENGINE POWER CONSUMPTION T.I.T.

/S PR KW(a) A g/J KG °K C0Nr1GuRATION (HP) (LB/HP—HR) (LB/SEC) (°F) BASELINE 422 78.2 1.30 1504 E. 11.3 AC 11.3 (565) (0.465) (2.86) (2250) REMOVE AXIALS 250 87.5 1.0 9.0 _C9 (335) (0.518) (2.20) (2040) C9 WITH IGV 197.7 91.9 0.79 1394 8.2 (265) AND REMATCH (0.544) (1.75) (2050) 32552 TABLE XLIII. RELATIVE COST SUMMARY (FOR EQUAL, WEIGHT) FOR DESIGN 3013-AC 11.3 COMMON CORE, DESIGN 3011-AC 11.5 FREE; TURBINE TURBOPROP AND DESIGN 3012 FREE TURBINE TURBOSHAFT.

COMPONENT RELATIVE COST, BASELINE TOTAL GEARBOX ^ COLD HOT ACCY & COMPRESSOR COMBUSTOR TURBINE A&T HSG HSG SYSTEM AIR INLET 0.20 0.16 0.06 0.20 0.02 (1,08 1 0.2 0.00 1.0 GATE 301: 0.24 0.16 n. 04 1 0.14 0.02 1 OD51 0.12 0.05_10.82 GATE 3013 WITH FABRICATION TECHNOLOGY 0.192 1 0.096 0.04 1 0.084 1 0.02 0.05 0.12 0.05 1 0.652 GATE 3011 0.04 0.20 0.25 0.16 0.02 1 Q 07 0.16 0.= 070. 97 GATE 3011 WITH FABRICATION TECHNOLOGY

0.20 0.096 j 0.04 10.12 0.02 Q 07 0.16

0.07 1 0.776 GATE 3012 w.^

- 0.16 0.02 0.07 5 0.70

0.04 0.20 0.16 0.0 GATE 3012 WITH FABRICATION TECHNOLOGY -- 1 0.096 2 O.n 4 0, 1 1 0.02 1 (10710.16 0.05 1 0.556

32553

TABLE XLIV. OEM COST: EXPRESSED IN DOLLARS FOR A PRODUCTION RATE OF 500 PER YEAR DESIGN CONSTANT WEIGHT CONSTANT POWER. CONSTANT P014ER NUMBER 227 KG (500 LB) 3£5.5 KW (490 HP) 422 KW (565 HP) 3013 81823 ---- 44885 3011 96791 "55891 ---- 3012 69849 33963 OEM PRICE ENGINE (a) COST A + B (b) RATIO K2 UNIT A (c) l GEARBOX A (d) 0.24390 0.8 1710.00 & AIR 2194.48 1710.00 I NLET S (e) 0.04878 0.8 484.48 A 0.09756 0.6 521.34 COMPRESSOR 1248.07 521.34 B 0.09758 0.6 726.73

T^l

COMBUSTOR A 0.04878 1.0 434.44 434.44 434.44 A 0.08536 0,6 456.17 TURBINE 1092.06 456.17 B 0.08536 635.89 0.6 A 0.01220 1.0 108.67 COLD 260.02 105.61 hlSG B 0.01220 1. o 151.41 A 0.03659 1.0 325.84 HOT 628.65 325.84 HSG B 0.02439 1.0 302.81 A 0.12195 110 1086.13 u AGCY 7386.94 1086.13 SYSTEM

B 0.02439 1. 0 302.81

A 0.04878 1.0 434.45 A&T 585.86 434.45 B 0.01220 1.D 151.41 TOTAL, OEM PRICE 7832.52 5076.98 f SELLING PRICE (1.5 X OEM) 11748.78 7615.47

32555

T (a) K2 : FABRICATION TECHNOLOGY FACTOR (b) AC 11.3 COMMON CORE (c) C9 (AC 11.3 DERIVATIVE) (d) A UNITS: PRODUCTION RATE - 15165/YR (e) 8 UNITS: PRODUCTION. RATE - 7525/YR ®r TABLE XLVI. RELATION OF THE COMMON CORE ENGINE PRICE TO THE PROJECTED FIXED WING AND HELICOPTER MARKET FIXED WING • HELICOPTER CATEGORY 11 111 U 111 P IV AG NUMBER OF ENGINES PER YEAR (a) 5480 1160 575 5940 1010 1120 (b) i WT.

DESCRIPTION KG ENGINE RETAIL PRICE $ (c) (LB) 93.5 - C9 ONE FRAME dK 8210 (206) C9 FRAME I 93.5 - 0E ------11445 (206) 65'8 C9 FRAME II 9450 (145) 92.2 AC 11.3 = 11750 (203) 78.1 C9 - DERIVED FROM AC 11.3 ' qc--7615 – – - (172) AC 11.3 FREE TURD WE 81.5 27350 (179)

P ERCEN T

FEATURE COMPONENT

COST

REDUCTION

ONE COMPRESSOR STAGE REPLACES 10 REDUCE L TWO NUMBER OF

COMPONENTS

ONE TURBINE STAGE REPLACES THREE 16

Tll^

VAPORIZING PLATE COMBUSTOR

2 NEW

REPLACES ATOMIZER DESIGN CONCEPTS FULL AUTHORITY ELECTRONIC CONTROL REPLACES HYDROMECHANICAL

CYCLE HIGH SPECIFIC OUTPUT REDUCES

BENEFITS ENGINE SIZE THEREBY REDUCES COST T',

TOTAL COST REDUCTION

49.1%

I

32065

I

REFERENCES: N-Fan for General Aviation" December 1973; Hamilton Stan- 1. ANON: NASA CR-114665 dard Division.

2. Anon: ""rcraft Piston Engine Exhaut Emissions Symposium", Lewis Research Center, September 14-15, 1976. NASA CP-2005 et al "An Upd ' ate of the Direct Injected 3. Jones, C., Lamping.D.

Stratified Charge Rotary' Combustion Engine Develop'inents at Cur- tiss-Wright". February 1977. SAE 770044 - 4. Anon; "Aircraft Engine Emissions A Conference held at Lewis Confer- Research Center, Cleveland, Ohio". May 18-19, 1977. NASA ence Publication 2021 5. "General Willis. , Edward A. t Aviation Energy-Conservation Research Programs at NASA-Lewis Research Center". NASA TM 73884^ October Emissions 6. Diehl, L. "Reduction of Aircraft Gas Turbine Pollutant NASA TM-78870 k Status Report". June 1978.

7. Anon: "Business Aircraft & FAR 36". January 1977. Business arid Commerical Aviation Subsonic Aircraft Noise 8. "Status Little, J., Russell, R.E. Report Reduction". Marcb 1976. ASME 76-GT-116 on the Design . of Air- 9. Rudey, R. "The Impact of Emission Statidards craft Gas Turbine Combustors". SAE 760909 10. Bresnahan, D., Sievers, G. NASA Quiet, Clean, General Aviation Turbofan (QCGAT) Program Status. December 1976. ASME 77-GT-77 11. The Status of Small, Cooled, Axial Flow Turbines. H.F. Due, A.E.

Easterling and J.E. Haas. In AGARD proceedings No. 229, "High Temperature Problems in Gas Turbine Engines. NATO, 19-23 Septem- ber 1977 12. Arnold, D. Balje, O.E. "High Temperature Potential of Uncooled Radial Turbines". ASME 77-GT-46 13. Wood, H., Performance Potential of Single-Stage Gas Turbine Engines. SAE 739135 14. Metzger, F.B., Worobel, R., "New Low-Pressure Ratio Fans for Quiet Business Aircraft Propulsion". SAE 730288 15. Federal Register, 17 July 1973, Vol 38 No. 136, Part II

} i -^ APPENDIX A LIST OF SYMBOLS AND ABBREVIATIONS Ag - Agricultural Assy - Assembly ATR - Area Taper Ratio B - British Thermal Unit CDP - Profile Drag Coefficient Cf - Coeffi cient CL - Lift Coefficient i f CO - Carbon Monoxide COMPON - Component Co - Isentropic Spouting Velocity CR - Cost Ratio f; dB - ° Decibels Del - Delivery Demo - Demonstrator Des - Design DIA - Diameter - DN Bearing Sore Diameter Times Speed DOC - Direct Operating Cost DOD - Department of Defense k DS - Directionally Solidified € E, Lt DS - Specific Diameter DTC - Design-To-Cost Eff'y - Efficiency Equivalent Horsepower EHP - EPA - Environmental Protection Agency EPNdB - Equivalent Perceived Noise Decibels 4 s

APPENDIX A (Continued) F Fahrenheit - OF Degrees Fahrenheit - FAA Federal Aviation Administration - Fab Fabrication - FAR36 Federal Aviation Regulation Part 36 - F/A Friel Air Ratio - FT Feet - gravitational constant, grams - LAMA General Aviation Manufacturers Association - H Enthalpy - Had Adiabatic Head - HP Horsepower - HR Hour - ICAO International Civil Aviation Organization - IGV Inlet Guide Vanes - IN Inch - Insp Inspection ;estrum Instrumentation - J Joules - JD Joules Constant - K Kelvin - 0 Degrees Kelvin K - KG Kilogram - KM Kilometer - Kilonewton Kn - Knots Nautical Miles Per Hour

APPENDIX A (Continued) - Knots, Air Speed KTAS LTO - Landing Take Off L C 2 - Limited Life Cycle Cost LB - Pound - M Meter - Materials for Advanced Turbine Engines MATE MAX - Maximum - Min Minute MISC - Miscellaneous M, m - Meter mm - Millimeter MPH - Miles p er Hour MTCE - Maintenance N - Newtons N - notational Speed Nautical Mile NM - - oxides of Nitrogen NOX NS - Specific Speed OBd - Objective Manufacturer OEM - Original Equipment ORC - Overrunning Clutch - Overall OV Pressure P - PAX - Passenger - Percent Cost Aircraft PGA - Powder Metal PM

(Continued) APPENDIX A PR -- Pressure Ratio PWR - Power QCGAT -- Quiet Clean General Aviation Turbofan OR - Degrees Rankine ' R/C - Rate of Climb RECIP - Reciprocating REL - Relative Req'd - Required REV - Revolution - ROI Return On Investment RPM - Revolutions Per Minute S - Second Sect - Section - SFC Specific Fuel Consumption SLS - Sea Level Static SLTO - Sea Level Take-Off S/N - Serial Number T - Temperature T1 - Engine Class Thrust Less Than 36.' KN (8000 1bs) TBO - Time Between Overhaul THC - Total Hydrocarbons Ti - Titanium - TIT Turbine Inlet Temperature Take-Off TO - - Take-Off Gross Weight TOGW

APPENDIX A (Continued) U - Wheel Speed V - Volume vs - Versus ==17.

VSTT - Variable Speed Training Target Wa = Airflow (Absolute) k WT - Weight

a

- Rate of Change of Column Heading

a Payload with Respect to Payload

a -

Rate of Change of Column Heading a sfc - with Respect to SFC p _ Incremental and/or Delta 1 1T AP - Pressure - Efficiency 7?

PO - Stagnation Density 11 - Micro - Y B Ratio of Station Temperature to Standard Temperature S - Ratio of Station Density to Standard Density 1,1-