QCGAT aircraft/engine design for reduced noise and emissions

QCGAT AIRCRAFT/ENGINE DESIGN FOR REDUCED NOISE AND EMISSIONS Leonard l'Anson and Kenneth M. Terrill Avco Lycoming Division INTRODUCTION The multi-engine general aviation fleet size is expected to increase by 70 percent in the decade of the 1980's according to the General Aviation Manufacturers Association (GAMA). These general aviation aircraft typically use suburban airports that are unprotected by commercial buffer zones. Con- sequently, there is the potential for general aviation to create more wide- spread adverse community reaction to noise and pollution than that experi- enced with commercial air carrier aircraft. Recognizing this, NASA let con- tracts to apply large engine quieting and emissions reduction technology to smaller engines and to develop new and more suitable technology where re- quired. These resulting "Quiet, Clean, General Aviation Turbofan" (QCGAT) contracts required delivery of a turbofan engine and nacelle demonstrator9 as well as preliminary defintion of an appropriate general aviation aircraft system_ that could use the engine as propulsion.

This paper describes the resulting aircraft/nacelle/engine designs created under the Avco Lycoming contract. These designs reflect the technical expertise of the following subcontractors: Aircraft Design - Beech Aircraft Corporation Nacelle Mechanical Design - Avco Aerostructures Nacelle Acoustic Treatment - Lockheed Aircraft Corporation AIRCRAFT PRELIMINARY DESIGN To guide the aircraft system design, five primary objectives were estab- lished: I. Practical, direct application of technology without significant scaling was very important. This required selection of aircraft and engine sizes which would be appropriate for an appreciable segment of general aviation.

. The aircraft must also offer attractive ranger fuel economy9 and flight speed. A target of 2593 kilometers (1400 nautical miles) was established. This exceeds the range of most current small business aircraft. It also provides non-stop travel between opposite extremes of high density traffic areas.

i01 o A cruise Mach Number of 0.62 was chosen as an optimum compromise be- tween time and fuel economy. It provides 40 percent higher cruise speed than a turboprop, with a 30 percent improvement in fuel econ- omy over operation at 0.8 Mach Number.

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A balanced field length of 762 meters (2500 feet) was desired be- cause it permits safe operation from 70 percent of all U.S. airports which are open to the public, including airports with sod runways.

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Ecological characteristics of an aircraft system are likely to be- come primary competitive parameters for general viation in the 1980 decade. Therefore, they deserve close attention in design selections.

The initial step in aircraft preliminary design was the selection of ap- propriate size and design. The vast majority of general aviation aircraft operating from airfields located in suburban communities are in the size class below 5433 kg (12,000 ib) gross weight. In the lower extremity of the gross weight spectrum, small private aircraft in the range below 1814 kg (4_000 Ib) are generally powered by single-piston engines. It is expected that market constraints for very low-cost aircraft in this class will dictate continued usage of piston engines for the foreseeable future. It9 therefore, follows that the greatest public ecological benefits can be realized by in- troduction of a quiet9 clean aircraft system in the 1814 - 5433 kg (4,000 - 129000 Ib) gross weight class. Figure i shoes the projected market volume for various sizes of general aviation aircraft.

As with the passenger car trend towards smaller and more sophisticated cars to perform the same function9 it is expected that the decade of 1980's will see a similar general aviation trend towards reduced aircraft weight and smaller engine size for the same mission. Because noise9 emissions, and fuel consumption reduce with engine size, subsequent improvement in ecological characteristics can be anticipated. Utilizing technologies such as turbofan propulsion, high aspect ratio super critical wing and lightweight composite structures, it is expected that a new class of small general aviation air- craft will emerge in the eighties. A target of 30 percent weight reduction was considered achievable.

For aircraft size selection, our target was the largest segment of gen- eral aviation aircraft where cost of turbofan propulsion does not preclude its introduction.

Figure 1 presents a composite plot of aircraft gross weight versus both "The Number of New Aircraft to be Built" and "The Current Estimated Nominal Aircraft Cost". The number of aircraft is based on General Aviation Manufac- turers Association data. The expected trend toward lighter weight and higher cost for the same mission has not been reflected to ensure conservative en- gine sizing. The range of 3175 - 4536 kg (7,000 to i0,000 Ib) gross weight appeared attractive, with 3629 kg (8,000 Ib) selected as our goal.

With the defined aircraft goals and Lycoming estimates for engine per- formance, Beech Aircraft Corporation conducted parametric studies to optimize the aircraft preliminary design.

The aircraft which evolved is depicted in figure 2. It is a sleek, ad-

vanced design, six-place aircraft with 3538 kg (7,800 Ib) maximumgross

weight. It offers a 2778 kilometer (1500 nautical mile) range with cruise

speed of 0.5 Mach Number and will take-off and land on the vast majority of

general aviation airfields. Advanced features include broad application of

composite materials and a supercritical wing design with winglets. Full-span

fowler flaps have been introduced to improve landing capability. Engines are

fuselage-mounted with inlets over the wing to provide shielding of fan noise

by the wing surfaces.

The high bypass ratio QCGAT engine plays an important role in shaping the

aircraft design. It offers a dramatic reduction in specific fuel consumption

compared with current pure jets and low-to-moderate bypass ratio turbofans.

Figure 3 provides this comparison, reflecting a 22 percent improvement in

fuel economy.

This lower fuel consumption may be used in either of two ways or in com-

bination:

It can substantially reduce aircraft gross weight for the same

range. The reduced weight provides compoundinterest on the fuel

economy. It also requires lower thrust favoring reduction of noise

and emissions.

If preferred, the lower fuel consumption can be translated into

longer range for the original gross weight.

Wechose to reduce gross weight and favor ecological characteristics.

Composite structures have been used extensively in the aircraft prelimin-

ary design to further reduce gross weight. Areas selected by Beech for the

application of composite materials are shown in figure 4. Kevlar graphite

composites were used for aircraft weight estimates. Further potential for

weight savings exists in the engine nacelles. Conventional design was used to

reflect the low-risk, low-cost test nacelle. Critical load carrying members

such as the wing spar are conventional aluminum construction.

Approximately 40 percent of the structure is fiber epoxy or honeycomb-

bonded structure. The use of composite structure in aircraft design provides

a decreasing rate of benefit as the application of composites becomes more

widespread in the design. Initial selection of applications is in noncritical

areas. As the stress in selected areas increases, the design safety factor

also increases to compensate for uncertainties resulting from the youth of

the composite application. Beech cautions that, while these composite appli-

cations are technically feasible, development beyond the scope normally

undertaken by industry would be required to assure success.

Structural design is in accordance with Federal Aviation Regulation Part

23 airworthiness standards for normal category airplanes.

A 17 percent thickness-to-chord ratio supercritical wing shape was selec-

ted because it offers a number of advantages over the conventional 12 percent

NACAshape. These advantages are summarizedin figure 5.

From the cross sectional comparison shown here, it can be concluded

that the supercritical wing provides larger volume for fuel storage

for the same chord width. The thickness increase has the supplemen-

tary benefit of higher section modulus, permitting lighter construc-

tion for equivalent bending loads.

The two shapes have comparable drag characteristics in the cruise

mode. Increase in the NACA airfoil thickness in an attempt to

achieve similar volume is impractical, because it results in a sig-

nificant reduction in useful flight speed combined with an overall

drag increase at lower speeds.

Iterative design studies show a 25 percent increase in fuel capacity

combined with a 3 percent decrease in aircraft gross weight. These

savings are for an equivalent aspect ratio of I0 and a design wing

loading of 2250 N/m 2 (47 Ib/sq ft) of wing area.

Prior test data have shown an appreciable increase in lift capa-

bility as depicted in this comparison. This promises a more forgiv-

ing aircraft for variations in angle of attack, enhancing safety.

For equivalent sophistication of flap systems, reduced landing

speeds are achievable resulting in shorter landing field length

capability.

The airfoil selected by Beech is similar to the NASAGA (W) - 1 airfoil9

but is tailored specifically for the high-speed, high fuel volume and the

high-lift requirements of the QCGAT configuration. The pressure distribution

used to guide the design tailoring would identify it as a BAC Sonic Plateau

airfoil with a 17 percent thickness ratio.

Full-span fowler flaps and spoilers have been introduced to achieve the

desired 762 meter (2500 feet) take-off field length and landing distances

with reduced wing area. Winglets have also been added to reduce actual span

and wing structural weight_ while maintaining high effective aspect ratio.

Major lift parameters are summarized below. Establishing optimum flap

settings was beyond the scope of this study. However, experience indicates

that a full flap deflection of 40 degrees for landing and take-off flap

setting of 40 percent of full deflection are appropriate for fowler flap de-

sign. These values of CLmaxrepresent available state-of-the-art with ad-

vanced airfoils.

C L d C L

CLmax

Flap Positions @ = 0 d

1.6

Up .132 .088

2.35

40%(Take-Off) .98 .088

3.45

Down 2.13 .088

Many drag influencing design details of the QCGAT airplane are not estab- lished at this time, because the airplane is as yet a preliminary design study. For drag analysis, ambitious estimates were made for the various items. Achievement of total airplane drag coefficients will require exacting effort in the practical development of the airplane. The resulting QCGAT air- craft drag compares with that of the Learjet Model 24, which is an extremely clean airplane. Allowances have been made for differences in wing thickness9 component sizes, etc. Drag coefficients used are summarized below: .02661, .02534 cruise Total CDp, Flaps and gear up .0164 Incremental CDp for landing gear .04066 Incremental CDp for full flap .0163 Incremental CDp for T.O. flap .01209 Incremental CDp for one engine out Four major airplane variables were considered in the parametric study to optimize the wing configuration. They are: i. Wing area 2. Wingaspect ratio 3. Fuel weight 4. Take-off weight In the study, for each performance goal, the limiting aspect ratio versus wing area is plotted for several take-off weights, including the effects of wing geometry on wing weight. These limits for each of the performance goals are then summarized on a graph so that the best compromise can be selected' A design point of 15.33 square meter (165 square feet) wing area and an effec- tive aspect ratio of I0 were selected.

Table I summarizes the expected weights for fuel, structure-plus-propul- sion, and complete aircraft with payload for both conventional and QCGAT air- craft designs.

The first line represents a hypothetical aircraft of current vintage de- sign with low bypass turbofan propulsion. Introduction of a QCGAT high bypass turbofan reduces fuel consumption by 22 percent. When this savings is iter- ated through the aircraft design, structure and gross weight reduce, provid- i05

ing an additional 5.5 percent in fuel economy. Similar iterations with weight

savings from composite materials and supercritical wing result in an addi-

tional 4.4% savings in fuel. The combination of engine and aircraft changes

provide 32%better fuel economy. The 22% reduction in gross weight permits.

the use of a smaller engine with 22% lower thrust and_ therefore_ lower abso-

lute emissions and noise.

Out aircraft study projected the maximumranges shown in figure 6 for

various payloads. While 1134 kg (2500 ib) is depicted as maximumpayload for

the aircraft_ only 753 kg (1660 Ib) is required to accommodate six people

with their baggage. At this payload 9 the achievable range is in excess of 2963 kilometers (1600 nautical miles). Flight conditions are 10058 meters (33_000 feet) and an average flight speed of approximately 0.5 Math Number.

In our QCGAT aircraft study_ landing distance_ rather than take-off capa- bility_ set the minimum usable airfield length. Introduction of full-span fowler flaps with moderate wing loading results in a very low "landing con- figuration" stall speed. The 32 meters/set (62 knots) stall speed compares with 41 - 46 meters/set (80 - 90 knots) for current typical jet and turbofan aircraft. Since landing distance is proportional to stall speed squared_ this low landing speed provides an attractive sea level FAR landing field length of 811 meters (2660 feet).

Figure 7 shows a representative sample of general aviation airfields plotted on coordinates of field elevation and field length. The Beech QCGAT aircraft with full useful payloads has a landing capability consistent with the majority of these fields.

The expected stall speeds promise a very forgiving airplane in the take- off and landing mode where most accidents occur.

The aircraft preliminary design was conducted to establish realistic cri- teria for noise measurement. Figure 8 depicts the locations for noise mea- surement_ as well as the aircraft and engine conditions at the point of measurement.

Approach noise is measured directly below the flight path 1852 meters (one nautical mile) prior to the beginning of the runway. Approach aircraft glide slope is fixed at three degrees. Take-off sideline consists of multiple measurements 463 meters (0.25 nautical miles) to one side of the take-off flight path. Take-off flyover condition is measured 6482 meters (3.5 nautical miles) from brake release_ directly below the flight path.

Looking at the tabulation of aircraft and engine conditions_ the approach conditions are quoted for 40-degree wing flap angle. This gives the shortest landing distance and the highest noise level. Where increase runway length is available_ 16 degrees flaps could be used. Velocity would increase to 55.6 meters/set (108 knots) and thrust would reduce to 818.47 n (184 Ib)/engine providing further reduction in noise.

Where take-off sideline noise is measured at multiple locations, the al-

titude of 262 m (860 ft) produces the highest estimated noise. Conditions are

summarizedfor this altitude.

Climb rate of the QCGAT aircraft approximates current aircraft of similar

mission. It attains an altitude of 106 m (3630 ft) at the take-off flyover

measurementpoint.

ENGINE DESIGN

This portion of the paper will touch on design objectives, noise, and

emission considerations, engine cycle and engine description, and conclude

with specific design features.

Before proceeding into the details of the engine design, a brief review

of the design objectives is in order.

The ecological characteristics of an aircraft system are a direct

reflection of the engine design. Careful attention to engine design

details which impact noise and emissions is required to produce an

engine that will becomea welcomeresident in a suburban community.

Appeal of turbofans is indisputable. They swept virtually the entire

commercial carrier market in a period of twenty years. The same

trend has started in general aviation with the larger size aircraft.

The rate of turbofan penetration into the smaller general aviation

aircraft size is a function of the engine cost. This cost generally

equates to simplicity of engine configuration. The result is basic;

to be successful, it must be simple in configuration. Take-off

thrust should be sufficient to permit operation from the majority of

general aviation airfields.

The mechanical design life goal should reflect the anticipated air-

craft mission. Beech projected a useful aircraft life of 12,000

hours with an average flight cycle lasting 90 minutes. Our design

goal was to match this life without replacement of major parts.

Despite the best intentions of the designer, parts do break and it

is desirable to be able to replace them conveniently. Modular engine

construction achieves this goal.

A 12.191-meter (40,000-foot) flight envelope is attractive for

avoidance of traffic and bad weather.

The need for fuel economygoes without saying.

The larger engines for commercial carrier aircraft have demonstrated sub- stantial advances in the technology of noise reduction. They have provided the recipe for quiet engine design which was used for QCGAT and is summarized in figure 9.

Blend low fan blade tip speed and low fan pressure ratio with high fan bypass ratio.

The fan stator should be set at least two fan blade chord lengths aft from the blade trailing edge. The quantity of fan stator vanes should exceed two times the number of fan blades to avoid inter- action of fan blade wakes with the stator vanes. Canted stator vanes are preferred.

Exhaust noise reduces with exhaust velocity, and turbine blade-pass frequency should exceed the audible range.

During the iterations which optimize an engine performance cycle, contin- uous attention is required to avoid adverse impact on emissions characteris- tics. Table II summarizes the primary causes for emissions along with the en- gine parameters which have a beneficial influence on emissions.

Unburned hydrocarbons and carbon monoxide emissions are primarily a reflection of poor combustor efficiency at idle. Low combustor inlet temperature at idle aggravates the carbon monoxide emissions. To re- duce these two constituents, one would strive for very high combus- tor efficiency at idle combined with elevated combustor inlet tem- perature. To achieve the higher inlet temperature, a compressor with poor efficiency at low speed if desired. This compressor should then be run as fast as idle thrust constraints will permit, and then bleed air to achieve even higher speed for the same thrust.

Whereas idle conditions have the primary influence on UHC and CO, take-off conditions predominate in the creation of NO X. Generally, the higher the combustor inlet temperature at take-off, the more difficult the problem with NO X. Another important axiom, NO X and CO can usually be traded through combustor design modification, Either emission can be improved at the expense of the other to achieve the desired combination.

Comments, so far, have ignored engine bypass ratio. Emissions are produced exclusively in the core engine. The higher the bypass ratio, the lower the emissions for a given thrust rating.

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A high-pressure compressor pressure ratio of approximately I0/i under

cruise conditions was selected as being achievable without compromise in con-

figuration simplicity. Demonstrated component technology indicated two tran-

sonic axial stages combined with a single centrifugal stage would be suffi-

cient. Modest work input requirements for this compressor permit selection of

a single-stage air-cooled turbine drive.

An NO X emission goal was considered the most difficult to achieve, with

high-pressure ratio engines requiring a complex combustor configuration. Be-

cause NO X emissions increase with pressure ratio, this i0/I pressure ratio

selection also favored combustor configuration simplicity.

Figure I0 shows the results of one of many parametric performance studies

conducted during the cycle selection phase. This particular study was conduc-

ted for 7620 meters (25,000 feet) cruise at 0.6 MachNumber. Engine perform-

ance is plotted two ways for comparison. The chart on the left provides bare

engine performance as it would be measured in an altitude test chamber. The

righthand chart modified the SFC coordinate to reflect installed specific

fuel consumption. Here9 losses associated with nacelle drag and weight are

factored in as the nacelle size varies with engine bypass ratio.

Performance for a variety of fan bypass ratios are plotted on coordinates

of specific fuel consumption and fan pressure ratio. In both figures, speci-

fic fuel consumption is seen to reduce with increasing fan bypass ratio up to

a ratio of i0/I. Optimum fan pressure ratio decreases with increasing bypass

ratio.

A fan bypass ratio of 9.6/1 at cruise was selected to limit required in-

put work to the capability of a single-stage fan drive turbine. The corres-

ponding fan pressure ratio was set at 1.35.

The engine was sized to produce in excess of 7117 N (1600 Ib) of thrust

under sea level static operating conditions. Sea level and altitude perform-

ance are summarized in table III. This performance compares favorably with

even larger, more sophisticated engines currently in use.

Figure Ii schematically shows the engine configuration we selected. The

gas generator section was not funded by the NASAQCGAT Program. The engine

has only six rotating cascades in total, and only two in the hot section

where maintenance costs normally accrue. As such, it achieves the simplicity

necessary to penetrate the mediumaircraft size general aviation cost barrier.

The configuration is a high-bypass turbofan composedof a single-stage

fan, a gas generator section, and a single-stage, low-pressure, fan-drive

turbine.

Initial compression is provided by the fan stage with the majority of the

air bypassing the gas generator section to produce thrust directly, much as a

small propeller would. Air flowing through the hub of the fan enters the gas

generator and is further compressedby the high-pressure compressor.

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A reverse-flow annular atomizing combustor accommodatedfuel burning and

energy release. Hot gases take a second 180 degree turn before flowing

through the high-pressure compressor-drive turbine.

These gases then continue axially aft through the fan-drive turbine.

Power from this turbine is transmitted forward by a shaft that is concentric

within the hollow gas generator shaft. Rotational speed is reduced by a re-

duction gear to match the optimum fan engine.

The advantages of the selected configuration are numerous. Initial stu-

dies projected attractive specific fuel consumption while maintaining the

desired simplicity of only six rotating cascades. Compliance with the recipe

for low noise and emissions has been achieved. The reverse-flow combustor

permits packaging the gas generator turbine inside the combustor to achieve a

short coupled engine_ thus avoiding difficulties of casing deformation and

shaft natural frequencies. The resulting engine center-of-gravity is close to

the axial plane of the main engine mounts, simplifying installation require-

ments.

Use of the reduction gear permits individual speed optimization to

achieve the best efficiency for the fan and the single-stage low-pressure

turbine. Also, the low-pressure turbine may then operate at higher rotational

speeds where blade pass frequencies, a common noise source, are outside the

audible tone spectrum even under low-speed aircraft approach conditions.

Overall_ engine configuration is shown cross sectionally in figure 12.

External dimensions are approximately 610 mmX 910 mm(2 feet X 3 feet)9 not

including the accessory gearbox.

Helical reduction gearing introduces an axial mechanical load which op-

poses aerodynamic loads on both the fan and the low-pressure turbine. This

provides a significant reduction in thrust loads on the ball bearings in both

the fan and low-pressure_turbine rotor systems.

The accessory gearbox is chin-mounted at the bottom of the main frame for

ease of maintenance without core cowl removal. Accessory drive is provided

from the high-pressure rotor spool via a conventional bevel gear mesh and

through-shaft with intermediate bearing support.

Figure 13 shows the modular maintenance features of the engine. The en-

gine disassembles into four basic modules, as shown. The fan module includes

fan, stators, reduction gear, and main engine frame. The core module contains

the high-pressure compressor, its drive turbine, and the combustor.

Separation of the gas generator and low-pressure turbine modules allows

visual inspection of all the hot-sectlon components. Full disassembly of both

modules was demonstrated to NASA representatives during the short period of a

coffee break at one of our coordination meetings.

On the wing maintenance is virtually unlimited by engine configuration.

Hot-section inspection, fan blade accessory replacements, gas generators and

low-pressure turbine module exchanges are but a few of the options available

to the operator.

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The QCGAT engine fan module is depicted in figure 14. The fan blade tip

diameter is approximately 559 mm (22 in.) with a modest tip speed of 335

meter/second (Ii00 feet/second) at take-off conditions. The ratio of stator

vanes to fan blades is 2.46 for acoustic considerations. The distance between

the fan blade trailing edge and the fan stator vane leading edge has been

maintained at 2.1 fan blade chord lengths to minimize noise from rotating

blade wakes. The fan stator is canted aft to maximize this distance for a

given engine length.

The reduction gear permits high turbine rotational speed producing a

blade-pass frequency which is above the audible range9 even under reduced

power approach conditions.

Fan blade containment capability has been provided in the fan shroud

design. Imbalance resulting from blade loss, has been a design criteria for

the supporting structure.

Hot oil sprayed into the hub of the spinner provides continuous anti-

icing and additional oil cooling.

Figure 15 shows an assembled fan wheel. There are 24 rugged long-chord

fan blades which are designed to withstand bird impact without the support of

a mid-span shroud. Fewer, long-chord blades were selected as being preferable

to a higher quantity of short-chord blades incorporating midspan dampers.

This reduces wheel cost and avoids the performance penalties associated with

midspan shrouds.

Figure 16 depicts the fan bypass stator assembly. In this design_ the

stator vanes are manually inserted into potted boots retained in the inner

and outer shrouds. This feature permits individual vane replacement rather

than returning the entire assembly for overhaul repair in the event of for-

eign object damage.

The QCGAT main structural frame is shown in figure 17. The frame is inte-

grally cast of aluminum. Four engine mounting bosses are provided to permit

selection by the airframe designer for top_ side_ or bottom engine mounting.

A cross-section of the QCGAT core engine is shown in figure 18. The com-

pressor and turbine stages are mounted on a hollow shaft which acts as a

throughbolt furnishing the necessary clamping force for the rotor system.

The compressor rotor thrust load is carried by a ball bearing at the

front end. A spring-loaded ball bearing at the end of the shaft permits ex-

pansion while maintaining radial position. Accessory drive is taken from this

rotor by meansof a bevel gear drive.

Two ball bearings, supporting the fan-drive turbine rotor9 are contained

in the samehousing which supports the aft high-pressure rotor bearing. This

avoids the cost of lubricating and sealing individual packages. A concentric

drive shaft through the high-pressure rotor delivers the fan-drive turbine

power to the fan module.

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All rotating cascades for both the high-compressor and low-pressure tur-

bine rotors are integrally cast to reduce cost. Air cooling is confined to

the high-pressure turbine.

Blade loss containment is provided throughout.

The combustor is a folded annular atomizing burner.

In a conventional atomizing combustor, an axial vortex is generated

around each atomizer by swirling the air with vanes. In the QCGAT size com-

bustor, sixteen atomizers and swirlers would have been required to attain

even, circumferential, temperature distribution.

Figure 19 shows the "circumferentially stirred" configuration which was

selected in preference to the conventional combustor. Primary air is admitted

through slots in the liner header producing flow circulation in a circumfer-

entially oriented vortex. Secondary air jets, called "folding jets", enter

the inner wall directly downstream of each atomizer and force the circumfer-

ential vortex into a horseshoe shape as it flows downstream. In this manner,

two downstreamvortexes are created for each atomizer, and the required num-

ber of atomizers is cut by one-half, to eight.

Prior testing has shown this configuration to be superior to conventional

combustors in emissions characteristics. It also demonstrated significant

margin in UHCand CObut was initially somewhatabove QCGAT NO X goals. This

permitted the trade-off of CO for NO X mentioned previously to assure

achievement of NO X goals.

As a result of this intensive design effort, the first QCGAT engine was

assembled in October 1978. In figure 20, the core engine module is shown

being connected to the fan module. Figure 21 is a front 3/4 view of the basic

engine fully assembled. The addition of the test inlet bellmouth, plus the

core engine cowling, is depicted in figure 22.

Figure 23 shows the birth of a new engine model installed in the test

cell just prior to its initial test run. Figure 24 is an enlarged view of the

engine installed. The first engine run was in October 1978. A 30-hour mech-

anical verification test was completed in April 1979. May and June were

devoted to damping an undesirable resonance in the ring gear. Emission tests

were conducted in July.

Figure 25 shows the engine at the acoustic test site during the acoustic

testing phase of the program, which was completed in August. The demonstrator

engine was then inspected, acceptance tested, and delivered to NASAin Octo-

ber 1979.

NACELLE DESIGN

A preliminary design in the flight nacelle was defined to establish a

realistic baseline from which a ground test nacelle could duplicate the im-

portant features at reduced program cost. Only the ground test nacelle was

fabricated.

An artist's conception of the flight nacelle is shown in figure 26. The

nacelle is composedof the following sections:

I. An inlet duct to provide uniform flow into the engine

.

A fan outer duct and core cowl to guide the bypass air around the engine .

A mixer assembly to force the mixing of hot, higher velocity core engine exhaust with the cooler, low velocity fan stream 4. A confluent mixing chamber preceding the final nacelle exit nozzle .

An aerodynamically shaped outer skin, designed to minimize drag at the higher flight speeds.

A mixed-flow confluent exhaust system was selected because it reduces the peak exit velocity, improves propulsive efficiency, and reduces jet noise.

Noise attenuation treatment in the form of perforated acoustic panels has been introduced in the air intake section and in the fan duct outer wall.

Figure 27 shows the engine/nacelle mounting and maintenance access panels. The engine is designed to carry the nacelle aerodynamic and "G" loads. An airframe or nacelle yoke attaches to two points on the engine main frame plus an aft steady link. The entire nacelle is then carried by the appropriate engine flanges. Four access panels are provided for ease of main- tenance.

The nacelle aerodynamic contours, summarized in figure 28, are optimized for low drag at 0.65 Mach Number, 10688-meter (35,000-foot) cruise condi- tions. The intake is designed for high cruise efficiency with modest compro- mises for static 20.6 meters/sec (40-knot) crosswind tolerance and pressure recovery at low-speed take-off conditions. A NASA/McDonnell-Douglas computer program for three dimensional flow calculations has been used for predicting inlet cowl flow conditions. This program computes flow over axisymmetric bod- ies at various flow angles of attack. Inlet and fan duct flow velocities are generally below 0.4 Mach Number.

The nacelle is circular in cross-section except for the bottom portion which expands into an elliptical section to house the engine accessories.

Boat tail angles vary from 14 to 18 degrees.

The QCGAT ground test nacelle which was used to explore noise and emis-

sions reduction is shown in figure 29. The internal flow lines, flight inlet

lip, and the exhaust nozzle are identical to the flight nacelle. External

skin was eliminated to reduce program cost.

Flight-worthy hardwall and noise attenuation panels can be readily ex-

changedas desired.

Three inlet lip configurations (figure 30) which were tested with the

QCGAT engine are as follows:

I. An inlet bellmouth for loss-free baseline calibration

2. An exact replica of the flight nacelle lip

3. A lip designed to simulate landing approach inflow conditions.

The flight lip and the inlet bellmouth are compared in figure 31.

Figure 32 shows two views of the mixer nozzle. Studies showedseven lobes

to be the optimum for our engine. We selected six, with a very minor perform-

ance penalty, to avoid any possible seventh-order upstream excitation of tur-

bine blading. Both shaker tests and engine strain gage testing showed satis-

factory dynamic characteristics.

CONCLUSIONS

Challenging objectives were set for the QCGAT aircraft preliminary design

to respond to our assessment of general aviation needs for the 1980 decade.

The aircraft design achieves these objectives to provide six-place, long-dis-

tance flight which will be attractive to both the user and the suburban com-

munity.

Flight characteristics of this aircraft have been computed to define

realistic criteria for measurementof ecological characteristics.

Reflecting on the engine and nacelle designs, the primary objectives of

the QCGAT Program have been fulfilled. Large engine noise reduction technol-

ogy has been successfully employed to the general aviation size engine. The

QCGAT Program culminated in demonstration of QCGAT acoustic goals with margin.

A simplified approach to emissions was conceived in response to this pro-

gram, and the QCGAT goals for emissions were very nearly achieved. Consider-

able margin was demonstrated for both CO and UHC emissions and NO X was

within I percent of the goal.

QCGAT has given birth to a new engine which is designed to serve the

needs of general aviation in the 1980's. While still in its infancy, it has

demonstrated attractive performance by current standards. Further development

tuning will be required to achieve its full potential which is reflected in

the QCGAT performance goals. Componenttests have verified the long term ob-

jectives. However, turbine rematch is required to recoup the inherent config-

uration performance.

Versatile ground test nacelles were created to investigate ecological

design parameters. Acoustic panels versus hardwall and mixed exhaust versus

split streams were tested under this program.

i15

TABLE I BENEFITS FROM ADVANCED AIRCRAFT DESIGN AIRCRAFT FUEL STRUCTURE AND GROSS WEIGHT, PROPULSION WEIGHT, CONFIGURATION Kg (LBS) WEIGHT, Kg (LBS) Kit (LBS) Current 1,368 2,518 4,518 Aircraft (3,016) (5,551) (9,960) Introduce 992 2,390 4,013 QCGAT Engine (2,186) (5,268) (8,848) Use Composites 931 1,978 3,538 and Supercritical Wing (2,053) (4,360) (7,800) 23% 22% SA VINGS I 32% TABLE II DESIGN CONSIDERATIONS FOR EMISSIONS EMISSION CAUSE Unburned Hydrocarbons Combustion Inefficiency Carbon Monoxide Inadequate: Residence Time, Temperature, Efficiency NOx High Residence Time/Temperature Smoke Local Rich Zones • Beneficial Engine Characteristics: High Combustion Efficiency at Idle High Combustor Inlet Temperature at Idle (High Speed, Low Compressor Efficiency, Bleed)

TABLEIII

PROPOSED QCGAT ENGINE PERFORMANCE FLIGHT CONDITION MAX CRUISE TAKEOFF 0.6 M, 7315 M PARAMETER Sea Level, Static (25,000 FT) 7206 N 2162 N Thrust (1622 Ibs) (486 Ibs) Speci,fic 0.0367 Kg/N/hr 0.063 (Kg/N hr) t-uel (0.360 Ib/Ib hr) (0.626 Ib/Ib hr) Consumption

NEW GENERAL AVIATION AIRCRAFT

FOR THE 1980 DECADE

NUMBER NOMINAL OF AIRCRAFT AIRCRAFT COST, $ x 106 -2 10,000.

5,000 - I I I I 12,000 (Ib) 4,000 6,000' 8,000 10,000 I I l I , I t 5,000 Kg 2,000 3,000 4,000 AIRCRAFT GROSS WEIGHT Figure 1

QCGAT AIRCRAFT DESIGN

TWIN FUSELAGE-MOUNTED _:i. QCGAT ENGINES ili : :i L_ i ......

..... LER FLAPS i :F_LL SPAN/ SUPER-CRITICAL WING :: !i!!::iii:_:_i_i_:ii:::'_'_ WlNGLETS WIDE USE OF COMPOSITE STRUCTURES Figure 2 -

PERFORMANCE

(Ib/Ib/hr) CURRENT AND FUTURE ENGINES Kg/N/hr 1.1 JETS AND LOW .11 - 1.0 .10 - 0.9 MODERATE .09 - BPR FANS O LL _0 0.8 .08 - QCGAT HIGHBPRFAN 0.7 _ADVANCED CYCLE .07 - I I I I I 0.6 .06 - - 100 90 80 70 60 50 40 PERCENT - MAXIMUM THRUST Figure 3

DESIGN AREAS SELECTED

FOR COMPOSITE STRUCTURES

KEVLAR SKIN ALUMINUM SKIN I[] GRAPHITE REINFORCED HONEYCOMB CORE Figure 4

WING SELECTION

17% 17% S.C.& CD NACA 12% NACA t i i - MN 0.5 0.6 0.7 0.8 NASA SUPERCRITICAL 17% CL _,CL MAX_ 17% S.C.

% NACA 35% £.)_w_z_ +1+200t_" _" " r' G_'1"_'11+25%1 GROSS WEIGHT _ -3°/° / n"w _10F i I f IX 0.12 0.17 THICKNESS/CHORD Figure 5

BEST RANGE vs PAYLOAD

I MACH NO. 0.5(MEAN) AT 11,480 M (35,000 FT)| MAX FUEL 975 Kg (2150 Ib) Kg (Ib)

I

3000, 1200 -

l

200O D 800!- < 0 Fu , < PLUS BAGGAGE I 400- t I I I I I I I I I _

00 I I I 1Woo

200 400 600 1000 1200 1400 1600 8OO (NM) I I I I I I I I I 0 400 800 1200 1600 2000 2400 2800 3200 Km RANGE Figure 6

AIRCRAFT STALL SPEED AND FIELD CAPABILITY

M/SEC (KCAS) SPECIFIC FIELDS SERVED BY LOCAL SERVICE AIRLINES IN U.S.A.

Km (FT) 3 10

40 I

o

QCGAT o 8 £3 f X LU LU 30 13_ I I I d - LANDII_G I--

_ 20

WEIGHT I

">' 4

i I I I I TAKEOFF

S 2

.

I WEIGHT W i.

,° I I I I i !

6 7 8 9 (Ib) 0 0 0 2 4 6 8 10 12 14(ft) I I I f I I I i I 1 2 3 4 M 4 Kg 0 WEIGHT x 1000 FIELD LENGTH x 1000 Figure 7

AIRCRAFT NOISE MEASUREMENT CONDITION

CONDITION ALTITUDE VELOCITY THRUST/ENGINE FLAPANGLE' M (FT) M/SEC (KTS) N (LB) 113 46.85 1157 Approach (370) (91) (260) 40 ° Takeoff 262 53.02 4880 Sideline (860) (103) (1097) 16 ° Takeoff 1106 53.02 4644 Flyover (3630) (103) (1044) 16 ° BRAKE APPROACH n;/FA,£F FLYOVER "-7"--- / JO 0 0 0 O 0 0 I TAKEOFF SIDELINE OF RUNWAY Figure 8

DESIGN CONSIDERATIONS FOR REDUCED NOISE

VANE NUMBER GREATER HIGH BYPASS RATIO THAN 2X BLADE NUMBER BLA GREATER THAN TWO FAN BLADE CHORD LENGTHS LOW EXHAUST VELOCITY LOW TIP SPEED AND PRESSURE RATIO NO INLET GUIDE VANES Figure 9

DESIGN STUDY

SAMPLE OF OPTIMIZATION TO DEFINE FAN PRESSURE RATIO CONSTANT COMPONENT EFFICIENCY CRUISE AT 0.6 MACH NO. 7315 M (25,000 FT) ENGINES SIZED FOR CRITICAL OEI CLIMB OUT AT 0.15 MACH NO.

I

, CYCLE PRESSURE RATIO 10.2 Kg/N/hr (Ib/Ib/hr) (Ib/Ib/hr) .07 .68 Kg/N/hr .72 8.6 OPTIMUM FAN .078 - BPR \ PRESSURE RATIO 8.6 .068 9.4 .67 9\ z< .076 - 9.4 .066 9.6 O LL .66 \ O0 9.6 .074 .70 .064 __:d_ .71 QCGAT .65 .072 - .69 QCGAT / .062 I _ '07i : I f I I 1.3 1.4 1.5 1.3 1.4 1.5 .06 - FAN PRESSURE RATIO Figure 10

SELECTED QCGAT

ENGINE CONFIGURATION

• Configuration Simplicity eSix Rotating Cascades • Superior Performance • "GOOD NEIGHBOR" Design Figure 11

QCGAT ENGINE CROSS SECTION

891.54 (35.1) 669.036 (26.3)

/

i DIA 302.26 (11.90) DIA NOTE: Basic Envelope Dimensions are in mm (inches) 482.6 (19.0)

ENGINE MODULES

QCGAT

FAN GAS GENERATOR LOW PRESSURE MODULE MODULE TURBINE MODULE I I I i / ACCESSORY DRIVE MODULE

J

Figure 13

QCGAT FAN MODULE

Figure 14 Figure 15

QCGAT REPLACEABLE VANE FAN STATOR

Figure 16

QCGAT ENGINE MAIN FRONT FRAME

Figure 17

QCGAT CORE ENGINE

Figure 18

QCGAT COMBUSTOR

MEAN VORTEX __ _._._ CIRCULATION PATH _)JC____ SECONDARY / . _'_ic

cm

FOLDING JETS _ ,"_-(_-T-

(0-'

VORTEX "_'__

CIRCULATION _ <_/_/:'4

PATTERN _ FUEL INJECTOR VIEW A--A

QCGAT ENGINE ASSEMBLY

Figure 20

QCGAT ENGINE

Figure 21

QCGAT ENGINE WITH BELLMOUTH

AND CORE COWL

Figure 22

QCGAT ENGINE IN DEVELOPMENT TEST CELL

Figure 23

QCGAT ENGINE AT ACOUSTIC TEST SITE

Figure 25

QCGAT ENGINE IN DEVELOPMENT TEST CELL

Figure 24

NASA - AVCO LYCOMING

Quiet Clean General Aviation Turbofan Figure 26

QCGAT NACELLE ACCESS PANELS

TOP EASY ACCESS PANEL Figure 27

QCGAT NACELLE EXTERNAL GEOMETRY

2289.81 (90.15) NACELLE MAX VERTICAL PLANE 533.4 ,.-4 _(21.0) 882.90 NACA 1 - 1(34.76) SERIES -i \ 377_A i CONTOUR i _ ..... _,_ -_ I _-6146.8 t_,+.oo) I ! / (242) ' J _RkDiUS 14o06, Q 528 32 (23.8) ' _ __ ___ --- _L_ DIA (2,.,0;8 AN = 1419.35 SQ.MM - 1 _'_ I 3°49'-_, _ (220 se. IN.) 4 °_ I j_____L___ -_ I JJ , ELLIPSE--_..._._.._ _ ' LT-_,_D"I_S v') 18°24J_ NACA-1 NACELLE WETTED AREA 5.11 SQ. M SERIES (55 SQ. FT) CONTOUR NOTE: Dimensions Are in MM (IN.) Excepted as Noted Figure 28

QCGAT FLIGHT NACELLE SIMULATION

::: ::: ::::: :: Figure 29

REPLACEABLE INLET LIPS

FLIGHT SYMETRICAL APPROACH NACELLE BELLMOUTH SIMULATOR Figure 30

FLIGHT NACELLE INLET

BELLMOUTH INLET

Figure 31

QCGAT MIXER ASSEMBLY

Figure 32