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NEW OPPORTUNITIES FOR FUTURE, SMALL, GENERAL- AVIATION TURBINE ENGINES (GATE) William C. Strack National Aeronautics and Space Administration Lewis Research Center SUMMARY The results of four independent contracted studies to explore the opportun- ities for future small turbine engines are summarized in a composite overview.
Candidate advanced technologies are screened, various cycles and staging ar- rangements are parametrically evaluated, and optimum conceptual engines are identified for a range of 300 to 600 hp applications. Engine improvements of 20 percent in SFC and 40 percent in engine cost were forecast using high-risk tech- nologies that could be technically demonstrated by 1988. The ensuing economic benefits are in the neighborhood of 20 to 30 percent for twin-engine aircraft currently powered by piston engines.
IN TRODUCTION The preceding portion of this conference was devoted entirely to business jet external noise and pollution, primarily, and fuel economy, secondarily.
This turbofan-powered segment of general aviation represents about one-fourth of U.S. engine net factory billings even though only 2 percent of the general- aviation aircraft engines sold are of this type (fig. i). The remainder of the conference will address the concerns of the other 98 percent. At the lower cost end are the single-engine airplanes powered by i00- to 300-hp piston engines.
About 23 000 of these were produced last year. The twin-piston powered air- planes utilize 200- to 400-hp engines of which about 8000 are produced annually.
The turboprop aircraft are mostly twins that require 500- to 1000-hp engines.
The U.S. produced about 750 of these last year, and Canada produced approxi- mately the same number. Collectively, these three categories represent 3/4 of the general-aviation engine net billings. Last year the total general-aviation engine billings were about 70 percent as large as those for the large commercial transport turbofans.
The principal problems facing these three categories are not so much envi- ronmental as they are economic, fuel, and safety related (fig. 2). Perhaps of greatest concern is the cost and availability of aviation fuels. Continued steep price hikes and the vulnerability of aviation gas to severe production cutbacks, or outright elimination, propels our quest for true multifuel power- plants. The safety of this class of aircraft continues to be questioned - with the spotlight alternating between airplane and automotive safety record compar- isons and the controversial one-engine-out twin problem. Engine dependability
is especially important in these smaller aircraft. Passenger comfort levels are
far less than those of larger turbofan aircraft, and the powerplant is a major
cause of the discomfort. Other concerns involve propulsion-related acquisition
and maintenance costs - especially for turboprop engines (fig. 3). NASA'sin-
volvement in addressing these concerns for the smaller general-aviation power-
plants is recent. Two years ago we recognized that the tiny amount of R&Tef-
fort devoted to small general-aviation turbines was not in proportion to their
actual importance. As the first step in rectifying that situation, we initiated
a series of analytic studies - known as the GATEstudies - to explore small tur-
bine technology opportunities. The question was: If we hypothesize a brand
new, small turbine engine that incorporates, say, 1988 level technology, what
size should it be; how should it be configured; and what benefits would it be-
stow upon us? The purpose, then, becameone of providing information to assist
us in planning future research. Wewanted to emphasize technologies that had
high payoff and high risk, but which could be ready for production development
by 1988 (given sufficient funding) and for mass production by the early 1990's.
These engines could be as much as i000 hp, but no more, but we emphasized sizes
below 600 hp since we perceived this size class to be potentially the most
rewarding - and challenging. Wealso emphasized aircraft cost of ownership as a
criterion of merit during the conceptual design process.
The first task of these studies was a 1988 market forecast that considered
all types of small airplanes and helicopters. This forecast determined engine
power sizes and other requirements of interest. Most of the study effort, how-
ever, was devoted to broad-scope parametric analyses wherein various cycles,
staging arrangements, and technologies were subjected to trade-off and screening
evaluations to determine optimum engine configurations for each importan t mis-
sion identified in the market forecast. Then, anticipating that the marketplace
could not afford different optimum engines for each application, an evaluation
was madeof a single common core to be used in a family of engines. And final-
ly, the required R&Tprogram was defined.
Weat NASA did someof these assessmentsourselves, but most were done un-
der contract to these four companiesworking independently: Garrett/
AiResearch, Detroit Diesel Allison, Teledyne CAE, and Williams Research. Within
a general framework, we permitted the companyteams the freedom to pursue direc-
tions and opportunities that they (rather than we) perceived as most attractive.
This freedom sometimes led to uniformity, for example, all companiesexpressed a
strong preference for turboprops instead of turbofans or turboshafts, and some-
times to interesting diversity, for example, the engine configurations and tech-
nologies varied considerably as did engine cost estimates.
The engine cost estimates were especially intriguing in the GATEstudies
because of the obvious opportunity to turbinize a portion of the piston-powered
market. The turbine engine is pretty muchaccepted as the most desirable type
of powerplant because of its many virtures - it has very low vibration levels,
high reliability, multifuel capability, a better safety record, low weight, few-
er emissions, less maintenance, and smaller installation losses. Despite these
advantages, the use of turbine engines has been blocked at about the 500 hp
level because of its higher fuel consumption and, especially, its 3:1 price pre-
mium (fig. 4). The challenge, of course, is to overcome the cost and fuel bar-
riers without sacrificing all the superior qualities. The trouble is that cur- rent technology does not allow this.
If we attempt to lower cost significantly, the efficiency suffers too much (fig. 5). But, if advanced technology could move the cost-efficiency band down far enough, we could certainly design a cost-effective small turbine or, if we choose, a better performer without cost reduction. On the other hand, if ad- vanced technology could not lower the band sufficiently, then only the high per- formance option is open. In the end, three GATE study team pursued the low-cost turbine versus piston theme, and the fourth pursued a high-performance, advanced turbine versus current turbine theme.
The results of the four contracted studies are presented herein by selec- tive examples that illustrate the main points in a representative fashion. The detailed results are documented in references i to 4.
CYCLES AND CONFIGURATIONS In figure 6, design turbine-inlet temperatures of 1800 ° , 1900 ° , and 2200 ° F are compared in terms of airplane total cost of ownership, fuel con- sumption, operating cost, acquisition costs, and engine cost. For all of these criteria, the optimum temperature level is 2200 ° F, or about 400 ° F above current small-engine levels. Although it appears that temperatures in excess of 2200 ° F would be even better, 2200 ° F was judged to be the highest temperature compatible with the materials available in the 1900's. The engine cost of the 2200 ° F engine is 40 percent less than that of the 1800 ° F engine because of a combination of factors. First, the physical size is about 40 percent smaller because the specific power improves substantially and because a smaller aircraft is required to do a given mission. In addition, the 2200 ° F engine incorpor- ates more cost-reducing technology, which retards the normal growth of cost with temperature and which keeps the cost per unit airflow nearly constant. Like- wise, the 2200 ° F engine weighs about 40 percent less than the 1800 ° F engine.
In fact, the airplane fuel consumption is improved 15 percent, not because the cycle efficiency improves (in fact, it is only i percent better), but because the engine weight is reduced. The engine weight and cost savings also produce 15 to 20 percent improvements in airplane acquisition cost, operating cost, and total cost of ownership° In a similar vein figure 7 displays cycle pressure ratio effects. In the lower plot, engine cost is displayed as a band that was drawn from four com- pressor point designs: a single-stage centrifugal at 9:1 pressure ratio and a relative cost of 1.0, a two-stage centrifugal at 20:1, an axicentrifugal at 11.3:1, and a three-stage axicentrifugal at 15:1. The band width indicates the increasing cost associated with more compressor stages at a fixed pressure ratio. Cost increases rapidly with pressure ratio as more compressor and tur- bine stages are required. Likewise, at any given horsepower level, weight in- creases too, so the power-to-weight ratio, shown in the upper part, becomes worse. Unfortunately, at the small airflows required in these applications (2 or 3 ib/sec), the cycle efficiency is not increasing rapidly enough to offset these adverse trends° In fact, as the SFC band shows, things are even worse
than that. While the 11.3:1 engine is 6 percent more efficient than the 9:1
engine, the 15:1 and 20:1 engines are actually slightly worse than the ii:i en-
gine because the componentefficiencies are suffering too muchat the very small
corrected airflows in the final stages. Hence, the minimumfuel solution is
about a 12 or 14:1 compressor pressure ratio, but the lowest aircraft cost of
ownership solution is about 9:1.
As a third example of these trade-offs, figure 8 presents a summaryof one
team's efforts to determine the best overall engine configuration for a medium
pressurized twin. They considered a turbofan with a gas-generator consisting of
a single centrifugal compressor hooked to a one-stage radial turbine. They also
considered a free-turbine turboprop in three different versions: The first has
the samesimple arrangement as the turbofan; the second has a two-stage centri-
fugal compressor; and the third has an axial turbine replacing the radial tur-
bine. Actually, the little diagrams only show the gas-generator portions, but
all four of these configurations also have a two-stage axial-power turbine. And
finally, they considered two arrangements of a single-shaft turboprop: both use
a single centrifugal compressor, but the first has one axial turbine following a
radial turbine, and the second has three axial stages. The evaluation criteria
are airplane total cost of ownership, fuel consumption, operating cost, acquisi-
tion cost, and engine cost, and the values quoted are all relative to the second
option - the simplest free-turbine turboprop. The bars are ordered from left to
right in the samesequence as the top little diagrams. The most obvious result
is that the turbofan is simply not in the running at all. Its penalties, which
are caused by its low efficiency at low flight speeds, run from 25 to 65 per-
cent. Actually, the optimum choice is not the simple free-turbine baseline but
rather the even simpler single-shaft configuration with one less turbine stage.
However, in consideration of other factors, especially commonality with heli-
copter turboshaft requirements, this team marginally preferred the free-turbine
baseline.
After manytrade-offs such as these and iterations with the marketing
analyses, the four teams settled on the cycles and configurations shown on the
right-hand side of figure 9. These engines are all turboprops ranging from 335
to 565 hp and are aimed primarily at the high-performance single-engine and
twin-engine airplane applications. For comparison, both a representative cur-
rent production turboprop (uncooled, old technology) and a hypothetical turbo-
prop incorporating currently available modern technology are illustrated on the
left-hand side of this figure. Allison's choice is a cooled, 2200o F maximum
turbine-inlet temperature, 14:1 pressure ratio, free turbine design. Two cen-
trifugal compressors are driven by two axial turbine stages, and another two-
stage turbine drives the propeller load on a second spool. This design differs
from both the current production engines and the hypothetical modern engines
mainly in having a better cycle and higher componentefficiencies. Its perfor-
manceis muchbetter, although its estimated cost differs little. Garrett also
chose a two-stage free power turbine, but selected a single 9:1 centrifugal com-
pressor driven by a one-stage radial turbine in their pursuit of lowering cost.
Teledyne's low-cost engine quest led to an engine family that is described
later. But its basic element is a very simple core engine which consists of a
single centrifugal compressor connected to a single radial turbine. This tur-
bine also drives the propeller load on a common shaft. Their turbine rotor is
uncooled and requires a very sporty, very advanced design, which is discussed
later. Finally, Williams Research sought low cost through a very unconventional
approach. Rather than the conventional idea of eliminating cost by eliminating
parts, they propose utilizing knownways of producing very inexpensive parts at
lower cost. The result was an uncooled, single-shaft, axicentrifugal arrange-
ment with six axial compressor stages and four axial turbine stages at modest
temperature but relatively high pressure ratio. This concept is described in
more detail later also.
ENGINE PERFORMANCE ANDCOST
The performance estimates for these engines are summarizedin figure i0.
Usually we think of SFCrising smoothly as we decrease engine size because of
adverse scaling effects. However, whenever a new engine is introduced, it may
distort our expected curve simply because of its advanced technology relative to
older, well-established engines. This happened a few years ago when the T700
was introduced. It yanked the curve down to form a "knee" in the trend curves
at 1500 hp. Exactly the samething would happen again if GATE technology en-
gines were introduced at 400 to 600 hp, since they would be 20 percent more ef-
ficient than current production engines of the samesize. The technologies that
lead to this are described later.
The three low-cost-theme study teams projected engine costs in two ways.
The first presumesno increase in production rates and simply reflects the in-
trinsic cost-reduction potential of using advanced technology (fig. ii). The
magnitude of this saving is about 40 percent. In other words, GATE engines
would be 40 percent cheaper to produce than today's engines. But, once a saving
of this size materializes, it would trigger increased sales, and this opens up
the possibility of a newmanufacturing facility dedicated specifically to GATE
engines, which, in turn, would cause even further savings - for a total reduc-
tion of as muchas 60 percent. At the sametime, market demandwould increase
to the neighborhood of I0 000 engines/year/company (assuming that two companies
split the market equally). Hence, without sacrificing too muchperformance, the
pursuers of the low-cost theme are predicting that GATEtechnology could provide
the key that unlocks this potential. To put the cost estimates in better per-
spective, figure 12 showsengine specific cost estimates for all four companies
against a backdrop of the current cost situation. Current turboprops cost about
three times as muchas piston engines. But remember,turboprop production rates
are two orders of magnitude less than piston engine rates. Allison's rather
sophisticated machine is estimated to cost about the sameas current turboprops,
and it triggers modest increases in sales. The low-cost theme estimates of
Williams, Garrett, and Teledyne at i0 000 units per year closely approach the
piston engine cost band, and this obviously represent_ a major departure from
today's scenario. Of course dollars per horsepower alone is not sufficient
since differing lapse rates, installation factors, fuel consumption, etc., are
equally important considerations. The net effect of all these factors is shown
later in the mission analysis results.
TECHNOLOGIES
But what are the technologies behind these improvements? Certainly nobody
could go out today and start building engines like these. Actually, it is some-
what difficult to succinctly summarize the advanced technologies identified in
these studies because each study team incorporated different ones - at least in
detail they are different. Nevertheless, figure 13 lists someof them in a com-
posite fashion, although not all of these would be present in a single design.
In the gearbox area the use of powdered-metal gears and laser hardening was
recommended by several teams to reduce cost and improve properties. Composite
material gear cases were recommended by Allison for stiffness and weight sav-
ings, while Teledyne suggested die-cast aluminum for a cost saving. Teledyne
also recommended a composite drive shaft, and nearly all teams recommended full
authority digital controls. However, the key elements in all of the concepts
involved the rotating machinery. Except for Williams, each team sought high-
performance centrifugal compressors using advanced analysis techniques, some
form of passive clearance control, and backward curvature. High stage loadings
without severe performance penalties were prevelant. And new manufacturing
processes, such as using powdered-metal titanium for the rotors, appeared.
Technologies for the core turbine were especially diverse. Two companies
selected high-temperature radial dersign: Teledyne, with an uncooled, powered-
metal concept and Garrett with a cooled, laminated construction process. Allison
selected a high-temperature axial arrangement with a cooled, dual-property rotor
and possibly ceramic stators. Again, passive clearance control was cost ef-
fective, and improvement of efficiency through better three-dimensional flow
analysis is required. Similar improvements were identified for combustors and
power turbines.
The first of several example key technology elements is illustrated in fig-
ure 14. This one represents the attainment of a 9:1 pressure ratio compressor,
at high efficiency, in a single stage. It requires advanced three-dimensional
blading with high tip speeds and high inducer Mach numbers, an improved three-
dimensional diffuser, better flow analysis and better experimental measurements,
improved surge margin, and a low-cost fabrication technique that yields essen-
tially a net shape part from powered-metal titanium. The benefit is to improve
compressor efficiency by 3-1/2 points relative to a current technology 9:1,
single-stage machined compressor, while reducing cost to be competitive with
cast designs. While the 4 percent, engine cost saving is not as large as for
someother components, it also saves 6 percent in engine weight, comparedwith a
two-stage compressor that would otherwise be required with current technology.
The second example is Teledyne's proposed uncooled, but high-temperature,
radial-inflow turbine. Its concept is based on their recent development expe-
rience with a 120-hp turbogenerator set for the Army plus someencouraging ana-
lytical work. Figure 15 shows the results of a preliminary analysis to verify
the concept's life potential. The stress-rupture life was evaluated for two
different blade geometries. One is relatively thin in the root region and has a
cross-sectional area that tapers down at the tip to 1/16 that at the root (i.e.,
it has an area taper ratio ATRof 16). The other is thicker at the root and has
a taper ratio of 31. The evaluation assumedthe use of equiaxed IN-100, a cur-
rent material, to give a high confidence level. Actually through, advanced
materials and directional solidification would probably be used to increase the
design's integrity. This design is very highly loaded, with tip speeds ap-
proaching 2500 ft/sec and transonic exit velocities at a maximum turbine gas
temperature of 2250 ° F. Under these conditions, the blade metal temperature
is 1800o F at the tips, and the lifetime is only 200 to i00 hr. However,
their engine is flat rated and will not require such high temperatures at take-
off or at any other normal condition. At cruise, the gas temperature is down to
1950o F, which yields 1552 ° F maximum metal temperatures. This yields a
3000-hr life for the 16:1 ATRdesign or a i0 000-hr life for the 31:1 ATRde-
sign. However, it is not certain that the 31:1 ATR is practical because of in-
creasing flow-path restrictions in the root regions (more detailed analyses are
required to determine an optimum ATR).
The third example is Garrett's cooled, radial turbine concept (fig. 16). It
consists of a set of photoetched laminates diffusion-bonded to form integral
cooling passages. After bonding, the part is electrochemically milled to the
full three-dimensional desired aerodynamic shape. Advancedpowdered-metal sheet
stock fabrication methods must be used to lower cost and thereby permit the use
of high-strength materials such as Astroloy. The net benefit would be a 9.8
percent efficiency improvement relative to current, cooled axial turbines while
reducing cooling bleed 20 percent and engine cost 21 percent.
The final example is that of an approach that does not apply specifically
to a single componentbut rather influences the entire engine. It is Williams'
unconventional approach to lowering cost through the use of restricted-geometry
blade and vane aerodynamic shapes. The concept is to design for very low,
rather than high, stress levels as depicted in figure 17. This allows perhaps a
150 ° to 200 ° F increase in turbine metal temperature without cooling, or,
with an advancedmaterial, muchhigher temperatures to exploit their specially
shaped temperature-stress curves. With MA 6000E, for example, an extra 300o
is possible. Either way, the lower stresses (perhaps 1/2 of conventional) imply
lower design speeds, and this, in turn, meanslower blade loadings, which permit
the use of low-cost, simplified blade manufacturing techniques. Specifically,
all compressor blades could have the sameairfoil section, be of constant chord
and camber, and be uniformly twisted; in fact, because only the lengths would
differ, the parts cost would be dramatically lowered. The corollary is that
higher pressure ratios are obtainable without muchcost penalty. Then all the
blades are held in place as the hub is formed around them in a single opera-
tion. The compressor vanes and all of the turbine airfoils are formed in the
sameway. The total result is a very different looking small engine concept
which attempts to achieve low cost without performance sacrifice by incorpor-
ating a large number of very low-cost parts instead of a very small numberof
relatively expensive parts.
COMMON CORE
Another concept for reducing engine cost involves using a common core for a
family of engines. Retaining parts commonality without sacrificing too much
performance is the key here because each of the diverse mission applications
prefers a different optimum engine. One approach to this dilemma is illustrated in figure 18 which shows Teledyne's C9 core engine slightly modified to accommo- date some additional parts that are required to reconfigure the engine for more power. This is done by adding a supercharging axial compressor stage, an axial turbine stage to provide the extra power, and a set of extra gears, which are duplicates of the first set to handle the increased power. This allows a 335-hp engine to grow 70 percent to a 565-hp derivative with on a 4-inch extension (from 34 to 38 in.), a 31-1b weight increase (from 172 to 203 ib), and a 54 per- cent increase in cost (fig. 19). At the same time, the SFC is i0 percent lower due to the increased cycle temperature and pressure and component rematching.
The price of commonality in this case is a 2 percent SFC penalty for the basic core engine. This results from the lower turbine temperature required to accom- modate a common fixed-area nozzle. The benefits of this approach to commonality are a 7 percent lower cost and a 16 percent weight reduction for the C9 335-hp version relative to the next best approach, which is using a single, large con- figuration and then shaving the flowpath area to reduce power.
AIRCRAFT BENEFITS The effect that these technologies, both individually and collectively, would have if GATE engines were installed in conventional but slightly improved airframes flying missions moderately more difficult than today's will be illus- trated with several examples. In each case the hypothetical aircraft is resized to accommodate the new engines. First, the Allison GATE engine was compared with a scaled turboprop version of their most recently improved 250 series tur- boshaft engine. Their GATE engine incorporates considerable materials and aero- thermodynamic improvement which accounts for higher cycle efficiency and smaller size. Specifically, a dual-property, axial, high-pressure turbine, while slightly more expensive initially, yields long life and much less engine main- tenance cost. Similarly, a transpiration cooled, Lamilloy combustor, while not inexpensive itself, allows the use of a short, compact, and long-TBO (time be- tween overhauls) combustion system. Ceramic rotors were not judged appropriate for manned aircraft application in this time frame, but Allison suggested that ceramic stators may be, although even they are only marginal. Lastly, a fiberglass/polyimide composite gearbox showed a slight cost advantage. Alli- son's advanced technology engine yielded 20 percent better SFC and 23 percent less weight. It costs 3 percent more to buy, but 35 percent less to maintain than a comparable current turboprop. A range of aircraft benefits are shown in the following list corresponding to the three aircraft types that they investi- gated (an unpressurized twin, a heavy twin, and a twin-engine helicoper): Technologies : i. Advanced materials and aerothermodynamics - higher cycle efficiency and smaller size 2. Dual property axial high-pressure turbine - much lower maintenance cos_ (5000 hr TBO) 3. Ceramic turbine stator - slight cost reduction 4. Lamilloy combustor - permits 5000 hr TBO at high temperature 5. Composite gearbox case - slight cost reduction Engine improvements : i. SFC - 20 percent 2. Weight - 23 percent 3. Cost -3 percent 4. Maintenance cost - 35 percent Aircraft benefits: i. Fuel burned - 23 to 32 percent less 2. Gross weight - ii to 21 percent less 3. Purchase price - +7 percent 4. Ownership cost -8 to 20 percent less Although the purchase prices do not change much, 23 to 32 percent less fuel is burned, and ownership costs drop 8 to 20 percent.
In a similar way, the other companies listed the technology elements that survived their screening processes and ordered them as shown in table I, a Garrett example. The benefits of each of the advanced technologies are given relative to a hypothetical, all-new engine using currently available technol- ogy. For example, the high-pressure laminated turbine technology raises the core turbine efficiency by 9.8 percent, reduces engine cost 21 percent, weight 7 percent, and SFC 7.4 percent and yields a benefit cost ratio of 561. The bene- fit is defined as the ownership saving over 20 years for a fleet of 15 000 medium-sized, twin-engined airplanes. The cost is the research investment re- quired to demonstrate technology readiness. The single-stage powdered metal titanium advanced compressor is i percent less efficient than a machined two- stage current technology compressor; yet it costs and weighs enough less to off- set this penalty. Another technology with large benefits is a low-pressure tur- bine that operates at a high work factor but low speed. Collectively, these technologies provide a 36 percent lower cost reduction, 20 percent lighter weight, and 13 percent better SFC relative to the best that we could do with today's available technology. One of the key elements is clearly the laminated turbine technology, which provides roughly one-half of the benefits.
Aircraft Fuel Now we can return to the most challenging issue, identified at the outset: comparing advanced GATE type engines with piston engines. One of the disadvan- tages of current turboprops is that they consume too much fuel: about i0 per- cent more than current piston engines for a typical twin-engine aircraft mis- sion. This is because their installed cruise thrust-SFC is inferior. GATE en- gines would eliminate most of this SFC difference as shown in figure 20. Since their installed engine weight is only 1/3 or i/4 as much as a reciprocating en- gine, the resulting GATE-powered airplane would actually save 5 to 15 percent fuel. Since avgas costs as much as 20 percent more per BTU, the real fuel cost savings are substantially greater than that.
Aircraft Economics A representative illustration of how a GATE-powered airplane compares with a reciprocating-powered airplane in economic terms is shown in table II for a light-twin airplane that cruises at I0 000 feet at 225 knots for ii00 nmi, is flown 500 hr/yr, and is sold after 3 years. The baseline is a current- technology, reciprocating-powered airplane that requires two 380-hp piston en- gines weighing 550 ib, each, that together burn 172 gallons of fuel. The air- plane takeoff weight is 6200 ib; the engines cost $ii 000 each; the airplane costs $207 000 total; it costs $51/hr to operate and, for the three-year owner- ship period, costs a total of $170 000. The percent changes for three different advanced engine options are shown in the right-hand columns. The first is an improved reciprocating engine presuming simply I0 percent lower SFC. It pro- duces rather modest aircraft economic improvements: 5 percent in total cost of ownership. Option 2 is a current technology turboprop, but produced at a rate of I0 000 units per year. It too is not very attractive - only a 3 percent net savings. Option 3, one of the low-cost GATE turboprops, is much more attrac- tive. This airplane would be 20 percent smaller and burn 8 percent less fuel, and, although the engine cost is up 23 percent, the complete aircraft cost is down 14 percent, and the operating cost is down 28 percent, for a total owner- ship saving of 20 percent.
If we expand our scope to include the other low-cost GATE versions and other applications (fig. 21), we see that, as a class, the twin-turboprop air- planes would cost 15 to 25 percent less to buy and 30 to 40 percent less to op- erate than their piston-powered counterparts. However, the benefits for high- performance, single-engine airplanes are only one-third to one-half as much.
Nevertheless, any economic benefits at all must be considered a bonus, inasmuch as the argument for turbinization could be predicated on noneconomic virtues alone. The obvious question is: When do these economic benefits disappear? A rough estimate of this is shown in figure 22 where a few data points from each study are plotted in terms of the reduction in ownership cost of GATE-powered airplanes relative to current reciprocating-powered airplanes as a function of the required shaft horsepower for the reciprocating aircraft version. The twin- engine airplane data looks impressive, showing 20 to 30 percent benefits. The single-engine data are too sparse to be certain, but it appears as though the economic incentive goes to zero somewhere in the 200-hp region. Of course, even at zero or slightly negative economic change turbinization is still attractive.
MARKET IMPACT Obviously, major benefits of this magnitude cause a large impact in the marketplace. The marketing forecasts that go along with the preceding are sum- marized in figure 23 for each study team in terms of the total number of turbine engines produced, both with and without an instantaneously mature GATE engine in 1988. Since GATE technology engines could not actually even enter service until the early 1990's, this is merely an indication of impact rather than an actual forecast. The picture is certainly striking because of the quite different es- timates. Allison's modest forecast is in agreement with their more conservative cost estimates, while Teledyne predicts a huge gain due to their lower cost es- timates and broad engine-size family. All of the estimates are much greater than the 1500 engines produced in 1976. Half of these were turboshaft engines for helicopters. However, the future GATE scenario forecasts that the turboprop would strongly dominate. A composite average of these four forecasts is shown in figure 24. A total of 20 000 GATE technology turbine engines would be man- ufactured annually, mostly turboprops, compared with one-fourth as many without GATE technology. The aircraft market results are shown with the pie charts, both with and without GATE technology engines. Without GATE, the turboprop share is forecast to grow from its current level of 2 percent to a level of 5 percent. With GATE, it would grow to about 35 percent, or a sevenfold in- crease. The twin-piston market would practically disappear, from 12 to 2 per- cent, while the single-engine piston portion would shrink from 68 to 47 percent, but it still would remain very large.
CONCLUSIONS A summary of what we perceive the major study result to be is displayed in figure 25. The most challenging, but rewarding, opportunity for small general- aviation turbine engines lies in the 300 to 600 hp region. Here, the proper combination of simpler design, improved materials, higher component efficien- cies, cheaper manufacturing technologies, and core commonality could result in sufficiently lower engine cost, SFC, and weight to overcome the traditional tur- bine engine cost barrier at the 500 hp size. Plotted here are the trends of aircraft cost versus engine size, and the large gap between reciprocating- powered and turboprop-powered aircraft is apparent. GATE technology permits large improvements in aircraft economies at the upper end of the reciprocating- powered class and fills in the gap between the relatively inexpensive recipro- cating aircraft and the expensive turboprop aircraft. In turn, this brings the many other virtues of turbine engines to a much broader spectrum of users and applications.
REFERENCES i. Baerst, C. F.; and Furst, D. G.: General Aviation Turbine Engine (GATE) Study Final Report. NASA CR-159482. (AiResearch-21-2997, AiResearch Manu- facturing Co. of Arizona; NASA Contract NAS3-20755.) NASA CR-159482, 1979.
2. Gill, J. C., et al.: Study of an Advanced General Aviation Turbine Engine (GATE). (EDR 9528, Detroit Diesel Allison; NASA Contract NAS3-20756.)
NASA CR-159558, 1979.
3. Smith, R.; and Benstein, E. H.: Advanced General Aviation Turbine Engine (GATE) Study. (CAE-1600, Teledyne CAE; NASA Contract NAS3-20757.) NASA CR-159624, 1979.
4. Lays, E. J. ; and Murray, G. L. : Advanced General Aviation Turbine Engine (GATE) Concepts. (WCR-78-113-15, Williams Research Corp., NASA Contract NAS3-20758.) NASA CR-159603, 1979.
Table I
ADVANCED TECHNOLOGY BENEFITS A 71, A COST, A WT, A SFC. BENEFffl TECHNOLOGY PTS % % % COST RATIO +9. 8 -21 -7 - 7. 4 561 HP LAMINATED TURBINE -L 0 - 4 -6 + L 4 232 PM Ti SINGLESTAGECOMPRESSOR ..... l 0 0 144 LOWCOSTFUELNOZZLES ..... 2 0 0 132 ELECTRONIC CONTROL HIGH WORK/LOW SPEEDLP TURBINE +6.0 - 5 -7.0 -7.0 498 ..... 3 O 0 226 LASERHARDENED GEARS ..... 36 -20 -1t O 402 (AVG) TOTAL NOTES L CHANGES ARE RELATIVE TO HYPOTHETICAL CURRENT TECHNOLOGY TURBINEENGINE 2. CLEARANCE CONTROL BENEFffSARE INCLUDEDIN ABOVE BENEFIT DEFINED AS OWNERSHIP COSTSAVINGSOVER20 yr FOR 15 000 MEDIUM 1WIN AIC 4. COSTIS NASA R &T COST SOURCE: GARRETT cs-79-42zo Table II GATE TURBOPROP AIRPLANES WOULD BE CHEAPER LIGHTTWIN AIRPLANE 1 CURRENT CHANGES TECHNOLOGY ADV TECH CURRENT GATE RECIP TURBOPROP 2 RECIP TECH TURBOPROP 2 (-10'kSFC) 38O -2 SHP. SLS TO - 11 -14 5.50Ib -3 ENGINE WEIGHT -68 -75 MISSION FUEL -Z0 10 -8 172gal 6 200 Ib -4 -15 GROSSWEIGHT -20 ENGINECOST $11 020 -2 113 23 ACQUISITIONCOST $ 207K -3 6 -14 OPERATING COST $ 51/hr -6 - 14 -28 -5 3 TOTALCOSTOF OWNERSHIP $ 170 K -20 ICRUISESAT 10 000 ft, 225 KNOTSFOR 1100n.m., 500 hrlyr FOR 3 yr 2ASSUMING10 000 ENGINESlyrPRODUCTION SOURCE: GARRETT CS-79-4190
U.S. CIVIL AIRPLANEENGINEPRODUCTION
1978, ESTIMATED _SINGLE ENGINEPISTON ...... l GENERAL _ MULTI-ENGINEPISTON AVIATION $516 MILLION 32 618
I TURBOPROPS
_ TURBOFANSIJETS
l
COMMERCIAL TURBOFANS $764 MILLION
I I I I I
0 4.00 800 8 16 24 RETAILBILLINGS THOUSANDSOF MILLION $ ENGINESSOLD C5-79-4192
Figure 1
G A PROPULSION CONCERNS
AVGAS l. Sr --'--JET A $1gal . 5 L '70 '75 '80 '8_ SAFETY/ • RELIABILITY • COMFORT __ _E-_'T YOU DO SOMETHING (f_,___._gg_ ,AX • .... ::.': .... (;,3-79-3713
Figure 2
• MAINTENANCE • COST CS-79-3712
Figure 3
CURRENT ENGINE SELECTION FOR LIGHT AIRPLANES
CS-79-I029 TURBINE DISADVANTAGES
Figure 4
EXPLOITINGENGINETECHNOLOGY DIFFERENTLY
1.0 RELATIVE
ENG,NEcosT I CURRE
ADVANC 0.5_ EJ
I
ol i i i
.1 .2 .3 .4 .5 cs-79-42os ENGINE EFFICIENCY
Figure 5
TURBOPROP MEDIUM PRESSURIZED TWIN
FREE TURBINE SINGLE CENTRIFUGAL COMPRESSOR TURBINE RADIAL GAS GENERATOR TURBINE 1. 5 TEMP.
1. 4 [] 2200o F [] 1900 ° F 1. 3 [] 18000 F RELATIVE 1.2 VALVE _ _ _ 1.]_,1.0 0.9 TOTAL FUEL OPERATING ACQUISITION ENGINE COST CONSUMPTION COST COST COST SOURCE: GARRET[ CS-79-4191
Figure 6
EFFECTSOF INCREASING CYCLE PRESSURE RATIO
400 SHP WEIGHT ........,....I--]| 1. 10 WEIGHT O.6 [-"
o.5
Iblhp I I. O0 O.4 i 2.0 rITURBINE AAA_///xr I_ _ _'_"_ RELATIVE 1. 5 COST YJ'J_"/_r'_:_-T UR BINE STAGE 1.0
--_c91 I I i I I
10 12 14 16 18 20 CYCLEPRESSURE RATIO SOURCE: TELEDYNE CAE c_-79-_,345
Figure 7
GAS GENERATOR CONFIGURATIONS
MEDIUM PRESSURIZED TWIN SINGLE-SHAFTTURBOPROP TURBOFAN FREE-TURBINE TURBOPROP
r--1 E3
tH
- _ I _ _I--t .ll _1
1. -- • 1. 45 I. 40 1.
RELATIVE VALUE 1,1 1, O _ -- 1 --4"_-- -
1 M
0.9 TOTAL FUEL OPERATING ACQUISITION ENGINE COST CONSUMPTION COST COST COST SOURCE: GARRETT cs-79-419_
Figure 8
FUTURETURBINEENGINEALTERNATIVES FOR SMALLGENERAL AVIATIONAIRC_-, T IT PIP CONFIG.
CURRENT PRODUCTI ON ALLI SON 22000 F 14 2C-2A/2A GARRET_ 18000 F 7-8 PIP i_(JJJJJJJJJ_ 2C - 3A 2C-AIA k-----_ _--"_'_---S_c-_ 22000 F 9 C - RI2A ETC.
TELEDYN____E CURRENT TECHNOLOGY " _1 22500 F 9 C - R 1850°F 12.8 6AC-4A 22000 F 8. 3 PIP -,,,_..,., _;._:_ _:: : 2C- 1A/2A TYPICAL) -_ L_ _.....
CS-79-3715
Figure 9
GATE SFC IMPROVEMENTS
.70 SEA LEVEL •66 -TPE 331 • 62 -C30 • 58 SFC 501D22A .50 TTOI .46 I I I ] .42 4 6810 0 1 2 SH P/IO00 CS-79-I023
Figure 10
BARRIERTECHNOLOGY
LOWCOST -- WITHOUT SACRIFICE IN PERFORMANCE APPROX. IO.O00 t PRODUCTION RATE: ENGINES/YEAR 1.0 40% 1 64% RELATIVE $/HP 0.5 CURRENT GATE GATE TECHNOLOGY TECHNOLOGY TECHNOLOGY ANDOEOICATEO MANU- CS-79-4204 FACTURING FACILITY
Figure 11
COSTREDUCTION FORECAST FOR GATE TECHNOLOGY ENGINES
A - ALLISON T - TELEDYNE G - GARRET]" W - WILLIAMS 150-- IO0 __ _ /COMPANY $/SHP (1977 OEM) 5O -- __1__ GATEAT 500/yr PiS__vr_Ta GATEAT 10 O00/yr
I I I I I I I I
%
200 300 400 500 600 700 800 900 CS-79-4206 THERMODYNAMIC SHP
Figure 12
GATETECHNOLOGIES
/-IQl-kx xA/1
/
cOMeR R --- CORETUR_,T--
GEARBOX • PM GEARS • LASER HARDENING • COMPOSITE CASE • DIE CAST AI CASE • CLEARANCE CONTROL • PM Ti ROTORS | • 3-D AERO
/
COMBUSTOR POWER TURBINE SHAFT • LOW-COST NOZZLES • HIGH WORK/LOW SPEED • COMPOSITE • VAPORIZING PLATE DESIGN • TRANSPIRATION COOLING I CONTROLS J • F.A. ELECTRONIC CS-79-4195
Figure 13
SINGLE STAGE CENTRIFUGAL COMPRESSOR
TECHNOLOG Y • HIGH STAGE LOADING • ADVANCED ANALYSIS • LDV MEASUREMENTS • OPTIMIZED DIFFUSION RATIO • 3D DIFFUSER • BOUNDARY LAYER BLEED • CLEARANCE CONTROL • GOOD SURGE MARGIN WlO VAR GEO • HIP PM TI NET SHAPE BENEFITS • +3.5 POINTS IN EFFICIENCY • 4% ENGINE COST REDUCTION • 69oENGINE WEIGHT REDUCTION CS-79-4214
Figure 14
RADIAL TURBINE LIFE
BLADE METAL TEMP, OF MAX CRUISE 1800 1552
_2
4- 1544 1390 1440 1298 1353 1218 RADIUS, in.
\/
/
ATR =16 ATR'31 STRESS RUPTURE LIFE, hr ATR = 16 ATR = 31 :3000 CRUISE (19500F) 10 OOOPLUS 20 MAX T.I.T. (22500 F) IO0 70 MAX -500 (22000F) 3OO MATERIAL: 1978 EQUIAXED IN-IOO SOURCE: TELEDYNE CAE CS-79-4194
Figure 15
INTEGRAL COOLED RADIALHP TURBINE
TECHNOLOGY • ADVANCED ANALYSIS • LDV MEASUREMENTS • END WALLEFFECTS • COOLINGFLOW OPTIMIZATION • MINIMUM iNCIDENCE LOSS • CLEARANCE CONTROL • LAMINATED CONSTRUCTION • ADVANCED ADB • ADVANCED ECM BENEFITS • +9.8 POINTS IN EFFICIENCY • 20 PERCENT REDUCTIONIN COOLINGFLOW • 21 PERCENT REDUCTION IN ENGINECOST SOURCE: GARRETT _S-79-4213
Figure 16
MANUFACTURING TECHNOLOGY AREASCOM PATIBLE
WITH RESTRICTED AERODYNAMIC SHAPES
LOW-COST S IMPLIFIED BLADE MANUFACTURE 4OK -- CONVENTIONAL HIGH STRESS 3OK-- DESIGN STRESS PS I 2OK -- \ / / LOW STRESS
1 " @
LOW SPEED / DESIGNS ,, / IOK -- / //
I I I I I I
1700 1800 19OO 2000 2100 2200 METALTEMPERATURE, OF • MULTIPLE ROW BLADE MANUFACTURE • BLADES IN PLACEAS HUB IS FORMED • ALL BLADES:. SAME AIRFOIL SECTION CONSTANTCHORD & CHAMBER UNIFORM 13NIST DIFFERENTLENGTHS CS-79-4203 SOURCE: WILLIAMS RESEARCH
Figure 17
ADD COMPONENTS TO C9 CORE
GEARING SUPERCHARGING COMPRESSOR STAGE CS-79-4202 SOURCE:TELEDYNE
Figure 18
COMMONCOREAPPROACH
CORE ENGINE (C9) -_ 34.O in.--_ l 1,35 SHP 0. 518 ESFC I I in.
2.2 Iblsec 172IbWEIGHTI 2040 o F
J
9.O PIP ADD SHADED COMPONENTS
+,>-
DERIVATIVE ENGINE(AC 11. 3)
565 SHP 0.463 ESFC 2.86 Ib/sec 203 Ib WEIGHT 2250 o F +54%COST 1L 3 PIP SOURCE: TELEDYNE CAE CS-79-4201
Figure 19
EFFICIENCY IMPROVEMENT PLUSWEIGHTADVANTAGE SAVES FUEL
TYPICAL TWIN AT 18 000 ft/225 KNOTS CURRENT • 5 TURBOPROPI INSTALLED 1. 5 RECIP ENGINE WEIGHT, INSTALLED , 4 r GATE CRUISE REC GATE Ib/SHP TSFC I I I •_ 200 400 600 0 200 400 600 SHP MISSION FUEL _-_\ _T _J CURRENTPISTON CURRENT TURBOPROP GATETURBOPROP BASE +10% -5 TO -15% CS-79-4200
Figure 20
BENEFITS RELATIVE TO CURRENT RECIPROCATING ENGINE
TURBOPROP POWEREDPRESSURIZED TWIN AND LIGHT 1M/IN ....... _.:.... ,;,.._,.._._ _, .,".._'.,;,_-:.!,_ :.: ::..; _,_:::'::,_,,;.;:'_:_,_.._i:.._;] :.'. t:.-",:_:'_z" TURBOPROP POWERED HEAVY AND LIGHT RETRACTABLESINGLE ENGINE 20 " 25_ 10 - 15% LESS GROSS WEIGHT I0- 15_ 5 - 1.5_ LESS FUELBURNED 15 - 25"_ 10 - 1591, LESS INITIAL COST 30 - 40'Io 7 - 1_o LESS OPERATINGCOST 25 - 35_ 8 - 15_/_ LESS LIFE-CYCLECOST HIGHER RELIABILITY GREATERSAFETYAND COMFORT QUIETERAND CLEANER MULTIFUELCAPABILITY CS-79-I031
Figure 21
GATE POWEREDAIRCRAFT HAVE LOWER COST OF OWNERSHIP
THAN EQUIVALENT RECIP POWEREDAIRCRAFT
ft knots n.m.
GARRETT [] MED. PRESS.TWIN 18 000 240 840 m LIGHTTWIN 10 000 225 1100 ELEDYNE 0 6-8 PLACETWIN 18 000 250 1200 @ 4-PLACE UTILITY 10 000 170 700 VILLIAMS /x 6-PLACEAEROSTAR 601TWIN (RETROFIT) • 4-PLACEMOONEY 201 (RETROFIT) 4C-- OWNERSHIP BENEFIT, 20 % GATE COSTOF 30- 100 200 300 400 500 cs-79-1o36 SHAFTHORSEPOWER FORRECIPVERSION
Figure 22
1988 CIVIL TURBINEENGINE MARKETUNDER 1000 SHP
TOTALOEM PLUS SPARES 40000-- 35 000-- r--1 WITHOUTGATE {SS3 WITH GATE 30000-- 25 000 -- TOTAL NUMBER OF U.S.
19 890 PRODUCED 20 000 TURBINE 16 000.
ENG INE S \\\\ \xxx ,\\"_] 15 000 \\\\ \\-.-_ . .\x| \\\\ x"x" " " x..\N ....\'._ \\\\ \\\-,i _AIRPLANES . .\\1 10000-- \\\\ \\\x,I \\\\ \\ \\\\ . .\\| \\\\ 611C \\\\ \\...
_6 _
477C_\_ 5000-- ,,-HELICOPTERS "--- _\\I / 1500 ,--, _'_"-.".1
' F':':I
•TELEDYNE 1976 WILLIAMS GARRETT CS-79-4207 ALLISON
Figure 23
1988 PRODUCTION FORECASTS
TURBINEENGINES GENERAL AVIATION AIRCRAFT( ~24 0OO) HELIC AIR_NES r .......TURBOPROP ...... -'\ / ', / //////////_ /' /_ -- -TWIN PISTON--- _ ', // ,/ ..._TURBOFANIJET-_ L _, x_ / " R HELICOPTE fAG "'7 w/o GATE
I
I
20 000 0 i0 000 w/o GATE WITH GATE CS-79-4193
Figure 24
GATE TECHNOLOGY COULD EXPAND DOMAIN OF
SMALL TURBINE ENGINES
IMPROVED MATERIALS / LOWER ENGINE COST : HIGHER COMPONENT PERFORMANCE_ LOWER ENGINE SFC i SIMPLER DESIGNS ti CHEAPER MFG. TECHNOLOGY ] LOWER ENGINE WEIGHT CORE COMMONALITY | EN'GINE COST BARRIER 1 000 000-- a _,r/,_, /- CURRENT " TURBOPROP _ POWERED AIRCRAFT PRICE, $ 116 O001_ G TE POTENTI L cs-79-1o35 0 200 4110 600 800 1000 ENGINE SHP