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From a presentation by Al Bosley in the 2012 SAE Congress Mobility History Committee technical session. This information is based on memories and historical material conserved by Robert F. Pauley; and an interview with Al Bosley by David Zatz.

The Chrysler Turboprop: Advanced Aviation Engine

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A portion of this was based on a short monograph by Robert F. Pauley, first published in the Journal of the Aircraft Engine Historical Society.

This is the history of an innovative aircraft engine concept, but even more than that, it's the story of Chrysler's engineering research section - a group of dedicated, talented scientists, engineers, designers, and support staff.

Starting in 1933, when it was directed by Carl Breer, Engineering Research was the advance "skunk works" of the company. The group was key in creating the Airflow cars, Oriflow shock absorbers, the XI-2220 aircraft engine, early Hemi engine concepts, the engine for the Chrysler Turbine Car (which was developed into the M1 tank engine), and a wide variety of other projects.

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Robert Pauley was part of the Research Department from 1953 until 1981, when he joined Williams International.

Al Bosley entered the Chrysler Institute in 1953, and later worked in Chassis engineering; was later appointed Director of Corporate Information Systems Planning. He returned to engineering for various administrative assignments prior to his retirement in 1991.

During the Second World War, Detroit automakers produced war materials; Chrysler had built tanks, trucks, antiaircraft guns, billions of rounds of ammunition, and parts for the atomic bomb.

In aviation, Chrysler had built 18,000 Wright R-3350 engines for B-29 bombers along with B-26 fuselage sections and other aircraft components.

In 1944, at Chrysler's headquarters in Highland Park, Michigan, the Research Department was wrapping up work on a major wartime development program, the 2,500 horsepower XI-2220 inverted V-16 aircraft engine for the Army Air Corps.

That program, directed by Chief Engineer George J. Huebner, Jr., began in May 1940; by July 1945, the engine was being flight tested in a Republic P-47H Thunderbolt - but the gas turbine engine was now known to be aviation's future. By 1945, the Army and the Navy were sponsoring the development of gas turbine engines, and the XI-2220 engine was canceled.

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These are many of the people worked on the V16 plane engine, including Carl Breer, near the center (in a suit); George Heubner; Dean Engle, later chief engineer at Dodge and later an engine man at Central Engineering; and Bill Chapman was in charge of testing the later turbine, later becoming head of Ford's turbine truck projects.

George Huebner, in his new position as Director of Research, had anticipated the switch to gas turbines for aircraft; and, early in 1944, he ordered preliminary studies for a company-funded turbo-shaft engine, based on their technical knowledge of thermodynamics and axial flow, or turbine, compressors. There was some discussion of a turbine engine for a Chrysler project, possibly as a truck engine; it's mentioned in some of the papers. For those purposes, they needed a compact regenerator and power takeoff via a reduction gear.

Creating the turboprop engine at Chrysler

The analytical work was assigned to engineers Sam Williams and John Jones, who had earlier designed a gear-driven axial-flow supercharger for the XI-2220 engine. They studied numerous engine configurations, all of which used a heat-exchanger to recover heat from the exhaust gasses to improve fuel economy, the major shortcoming of early gas turbine engines.

Regenerative or recuperative designs are a hot topic in gas turbines in 2017; the California Air Resources Board and others have been financing work on recuperative gas turbines which run on natural gas - as auxiliary power supplies that can come online very quickly when the grid gets into trouble. They hope that these turbines will be more stable, cheaper to run, and longer-lived with less maintenance than diesels.

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At the same time, other engineers in Chrysler's research laboratory were finding the combustion and metallurgical needs. As the XI-2220 program wound down, more staff started to work on the turbine program, which by mid-1944 had developed into a full-fledged research project and assigned a project number "A-86".

The technically correct practice is to use the word "recuperator" for a stationary heat-exchanger, and "regenerator" for a rotating type. The XT36 had a stationary heat-exchanger, but Chrysler's manuals and reports from the time use the term "regenerator;" so that is the word we will use for the heat exchanger.

In late 1944, members of the U.S. Navy Bureau of Aeronautics' Power Plant Division met with Chrysler executives, and asked if it would consider adapting this power plant for the Navy. They followed up with a letter, dated May 12, 1945, spelling out the requirements for a recuperated gas turbine engine (including a propeller drive, 3:1 compression-ratio axial flow compressor, tube bundle regenerator, and single-stage turbine).

Chrysler agreed to accept the Navy challenge. Work began officially on the project on July 1, 1945 after Chrysler had received a letter of intent for the contract.

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In October 1945, representatives from the Bureau of Aeronautics visited Chrysler. After expressing satisfaction with the progress, the Navy raised questions about the objectives in the original contract, which had been based on Chrysler's design studies using a 3.27:1 axial flow compressor, single-pass regenerator, and a single-stage turbine with an inlet temperature of 1,500° F. On November 20, the Navy informed Chrysler that they wanted an engine with more power and better part-load fuel economy, and the contract was almost completely rewritten to reflect those new goals.

The engine was now named XT36-D2 and Chrysler project number A-86A. The new contract called a 6:1 pressure ratio, a more efficient regenerator, and a two-stage turbine with an inlet temperature of 1,600° F.

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The contract covered two years, but it was understood that only a portion of the work could be completed by then; four later amendments extended the contract through June 30, 1949. The contract also emphasized the classified nature of the program; Chrysler was to develop the engine concept on its own, without exchanging information with other gas turbine manufacturers.

Bob Pauley was the primary designer; he made these drawings, and took the pieces and specs that the engineers (Sam Williams and others) put together - no computers or CAD. They were created full-size on a drafting board; I reduced them for this presentation.

The project ended up running from mid-1944 to mid-1948, four years; it was not a fast project, and it took a long time to get going. There were a lot of discussions.

Creating the XT36-D2 renegerative turbine engine

The outstanding feature of the XT36-D2, which made it unique in its class, was making the regenerator an integral part of the design. The primary emphasis was on part load fuel economy, targeting specific fuel consumption of 0.434 pounds of fuel per horsepower per hour at 30,000 feet at 60% cruise. The specified sea level rating of 1,000 horsepower at takeoff was of secondary importance.

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There were many studies to find the most compact engine package. A tandem arrangement would have been most aerodynamic, but would create an engine that was far too long. Putting components in parallel kept the distance between the compressor/turbine shaft bearings short; and, by putting the regenerator on the outside of the engine, the compressor and turbine housings would be the engine's "backbone." On January 5, 1946, study #13 was selected.

The heart of the engine was a single-shaft axial-flow compressor, more efficient than a centrifugal compressor, and able to reach higher pressures. It was driven by a two-stage axial turbine with a maximum inlet temperature of 1,600° F.

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This was the final layout, in blue. And you can see, it picks up from a small inlet into the inlet chamber and goes down the axial flow turbine - single stage. It makes the turn and goes through the heat provided by the exhaust. It goes into the combustion chamber and into the turbine at the right end of the drawing. It's made a full 180 degree turn, now flowing back; it's actually made lots of 180 degree turns, so the exhaust back pressure was pretty high. That's one of the issues in gas turbines and jet engines - losses in the exhaust.

The major layout drawings were finished by mid-1946, and by the end of that year, drawings for individual parts were being sent to the shop, and the lab began running tests on individual components.

In March 1947 a large, 2.5-size, four-stage compressor test rig was built to test blade airfoils; the rig was sized to match an existing dynamometer, but using the larger compressor also made it easier to test. Small compressors have to run at pretty high speeds for testing; increasing the size lets you get some of the same effects at lower rotational speeds. This was only a four-stage design; and the one in the "proposed production engine" was longer.

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Since it ran at the same Reynolds number as the proposed engine, the results were directly applicable to the final design. A burner test rig was also built to evaluate burner designs, and a full-size regenerator test rig was built to test that unit. Fabricating a leak-proof three-pass regenerator was challenging; the test in that rig did not run until October 1947.

The compressor was an axial-flow design with an steel rotor drum that had a constant diameter of 9.50" at the blade root and a wall thickness of 0.20". The compressor rotor had an inducer stage followed by fifteen blade rows and an exducer to convert the flow to a pure vortex. The inner wall of the stator housing, which formed the structural backbone of the engine, was tapered towards the rear to maintain a constant axial flow.

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Why did they make a half-size wood mockup? Partly to help explain the concepts to non-engineers, who don't visualize these things as well; partly to verify the packaging.

When I was in chassis engineering, we had a very large prototype mockup facility, which did the engine compartments for all new vehicles. I estimate 90% of the job you do as a chassis engineer (at the time I was there) was not designing parts - it was "designing" clearances. Could you assemble, get the parts in and out? Could you get the tools on them to work on them? Would they interfere with other parts as they moved? Manually cutting all the sections on a drafting board would take a huge amount of time - not like today; the computer does it all in seconds.

The blades for the inducer stage and the first three symmetrical stages were machined from magnesium alloy. Blade rows five through twelve were machined from cerium magnesium alloy and the last five stages were machined from 416 stainless steel. At assembly, their threaded shafts were inserted through holes in the compressor drum, or in the stator housing, and held in place with a single nut and washer. Blade angles were set using a gauge and blades were locked in place with glue and by staking the nuts. The compressor rotor assembly was coated to resist abrasion and corrosion.

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The compressor rotor/turbine wheel assembly was supported by two straight roller bearings spaced 43" apart. The bearing chambers were sealed with carbon rings and pressurized with bleed air from the compressor. Shaft speed was 16,900 rpm at military rating with 120% (20,300 rpm) used for maximum design over speed calculations. A tapered-land thrust bearing carried a combined axial load of over 10,000 lbs from the compressor and the turbines.

One major issue is getting those things in balance. They're rotating at several times, sometimes as much as ten times, the rotating speed of a reciprocating engine. The balance of the rotating parts, let alone those that in some cases have to reciprocate (like pumps), is pretty tough to achieve.

There is another problem with turbines, which these guys probably didn't have time to deal with before the engine was cancelled.

In the military engines that I worked on at Allison (as a summer intern in 1952-53), a jet engine is typically supported by front and rear engine mounts which are attached to the housing (or case), with the rotating parts inside. Inside a military jet, where you do high-G maneuvers, all of those parts can have up to 10 Gs applied to them, so the engine slightly bends.

Simultaneously, those loads are applied on the rotating assembly, and it bends. And if the rotating assembly and the casing do not bend through the same arc, the blades will drag on the edge and then you'll lose a blade and the thing will swallow itself. People get very angry about that, particularly the pilots.

This was one of the more complicated problems I worked on at Allison, trying to calculate the rotating stiffness of the J71 turbine compressor module to match the bending axis of the engine housing under G-loads. We were doing most of this on mechanical calculators; we had one very, very early and primitive digital computer.

There were some pretty good guys there who completed this work!

The regenerator was a three-pass, cross-flow design in the shape of a hollow cylinder with an inside diameter of 26.00", an outside diameter of 36.17" and 20.25" long. Early in the program, Chrysler had contacted two vendors with experience in the design of heat exchangers - the Harrison Radiator Division of General Motors and Trane - but both were heavier and less efficient than Chrysler's tube-bundle design.

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Chrysler's design had 6,927 Inconel tubes copper-brazed to 0.140" thick Z-nickel end plates at the front and rear walls of the regenerator. The tubes had an outside diameter of 0.219". The outermost bolt flange on the regenerator housing was subjected to a load of 30,000 pounds due to the unit's internal pressure. All of the tubes had spacer rings to prevent blockage of the gas passages in the event of tube bending.

Discharge air from the compressor (at 600° F and 95 psi) was ducted to the front chamber of the regenerator, and then radially outward to pass aft through the outermost tube bank, which contained 2,010 tubes. The air then made a 180° turn to pass through the central tube bank, which contained 2,352 tubes, and then another 180° turn to pass aft through the inner tube bank, which contained 2,565 tubes. On each pass, the compressed air in the tubes picked up heat from the hot exhaust gasses that flowed out through the regenerator and exited the engine through stainless steel exhaust outlets. These exhaust ducts were pointed aft, to give the engine an additional 217 lbs of thrust.

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This is the heat exchanger; you can see the three different passes in which the exhaust gases are going out and the compressed air is going through. The vertical burner tubes were fairly conventional at that time.

The regenerator weighed 175.5 lbs, had a total heat transfer area of 647 square feet, and at maximum engine speed delivered air to the burner chambers at 1,200° F. Regenerator effectiveness ranged from 78% at idle to 68% at takeoff. A sample three-pass tube-bundle regenerator ran for over 500 hours in the test rig with no cracking or leakage problems.

The burner had ten individual tubular chambers around the turbine housing, with a burner tube inside each chamber. The fuel nozzles and igniter plugs were near the rear, easily accessible for service.

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Chrysler tested eighteen different fuel nozzles from various vendors, finally selecting their own return-flow nozzle. The Type 310 sieve-type stainless steel burner tubes had holes for primary and secondary air. Concentric-cone-type burners were also tested, some made of Inconel and others of stainless steel, in a small-scale burner test rig. Heated air from the regenerator entered the burners at 1,200° F and, after mixing with fuel and burning, the hot gasses were directed into the first stage turbine nozzle at a maximum temperature limited to 1,600° F and at a pressure of 86.5 psi at sea level static.

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From what I understand from Bob Pauley, the burner cans were one of the major problems in development, as they were in all jet engines at this time. The air flows through the burner can. Fuel is injected, and like there's a spark plug. If it works right, it's supposed burn continually, without having to fire the spark plug continuously. But in the early days most had to fire the plug continuously, and the spark plugs had to be changed after every flight. They worked at least, half a dozen different alternative burner designs. There was one alternative, shown on the right, that they were starting to develop, I think, when the project came to an end.

The two-stage turbine had forged discs with welded-on buckets and integral shrouds to reduce blade vibration. The first stage wheel was 16.34" in diameter while the second stage was 18.12". At steady running conditions, the first stage turbine blades saw temperatures of 1,425° F and the second stage blade temperatures were 1,070° F. A radiation shield was placed on the aft side of the first stage turbine, and cooled with compressor bleed air. An inter-stage diaphragm between the two turbine wheels was supported by the second stage nozzle block.

The gear box, a three-piece magnesium casting, was at the front of the engine. It housed dual planetary gear sets that reduced the high speed (16,900 rpm) compressor/turbine shaft to a 1,200 rpm propeller shaft speed. The gear box also supported the propeller shaft and had mounting pads for accessories (two generators, a tachometer, and fuel pumps); the central housing also included lateral air intake ducts for the compressor.

Planetary gears were selected to save weight; the high speed gears had a reduction ratio of 3.75 to 1, with five planet gears instead of the more conventional three gears, which tend to equalize the loads. The low speed planetary also used five planet gears with the same ratio as the high-speed set, for an overall ratio of 14.06 to 1. Using five planetary gears required extremely close machining tolerances, but was necessary to transmit 1,450 horsepower from the compressor/turbine shaft.

As mentioned before, by mid-1946 the major layout for the XT36-D2 had been approved by the Navy's Bureau of Aeronautics, and Chrysler was given the OK to make detail drawings and fabricate parts for four engines plus spares. By the end of 1946, the Research Section's engineering and drafting staff had completed the major design work on the engine and were then free to work on improvements to the original design.

The second turboprop design: XT36-D2

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As a result Chrysler, negotiated an addition to the contract to develop an upgraded version of the XT36-D2 to be known as the A-86B (Navy designation XT36-D4). The goal was to go up to 3,800 horsepower, improve the fuel economy, and increase the life of critical hot components; performance increases were to be done by increasing the cycle temperature, requiring a complete redesign of the aft end of the engine. The initial goal was to increase the turbine inlet temperature from the existing value of 1,600° F to 2,000° F, and to ultimately to reach 2,400° F.

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Above, on the left, you can see the original two-stage turbine; they with shrouds at the outer ends of the blades. One of the energy robbers in gas turbines is hot gas leaking over the ends of the turbine blades, that energy doesn't drive the turbine to produce power.

How close can you get the turbine blades to the case, without rubbing? Everything changes temperature from the time the engine is started, so you don't want to rub them when you're starting the engine (perhaps in Alaska) from cold. It gets up to 1,700° Fahrenheit - perhaps. Another issue, how do you keep them cool? They didn't have the materials that we have now, so they investigated a system in which air was pulled up through the hollow blades, as you see on the right-side picture.

Sam Williams was in charge of that program, and he devised a novel turbine and nozzle cooling arrangement that allowed compressor bleed air to flow through hollow nozzle vanes and turbine blades. In addition, the individual burner tube chambers were replaced by a cylindrical cover with all ten burner tubes located inside an annular chamber.

The new gear box also included redesigned air inlet ducts to increase airflow into the compressor. Parts for the new "hot section" were on order in March 1948, and Chrysler was planning to run the A-86B version by the end of 1949. However, when the contract ended in mid-1949, those plans came to an end; and the A-86B design was never run or tested in an engine.

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By March 1947, parts for the XT36-D2 were being delivered and the first engine was slowly being assembled. At the same time test engineer Bill Chapman and the lab technicians were adapting Chrysler's 2,800 horsepower dynamometer to fit the new turboprop engine; that was the same dyno was used in the XI-2220 engine.

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The first XT36-D2 engine to be assembled, Serial Number 1, was run on this dynamometer for the first time on June 30, 1947 - just over 20 months since the revised contract had been negotiated. All of the early test engines had dummy regenerators so that basic engine performance could be evaluated, without the heat exchanger.

Shortly after that first startup, test engineer Bill Chapman noticed a steep temperature rise in the compressor thrust bearing and immediately shut the engine down. Several more startups showed the same unexpected overheating problem. That same thrust bearing had been thoroughly tested and developed in a test rig and had proven that it could easily take the high thrust loads. After disassembly the reason for the overheating problem became obvious-the test rig rotated clockwise whereas the engine rotated in a counterclockwise direction! A new thrust bearing was made, with its tapered lands correctly oriented, and there were no more problems.

The engine was run for 4 hours and 37 minutes when it suffered a catastrophic compressor failure at 12,000 rpm. Some of the compressor blade attachment nuts had failed, so positive locking rings were added to the vanes on a new compressor and a new stator. With those new parts, the engine was run for 13 hours and 39 minutes and then removed for a complete inspection.

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They used the same test stand, equipment, and such that they had used in developing the XI-2220. The Navy paid for all the conversion costs and necessary equipment, as part of their navy contract.

Meanwhile, Engine Number 2 was installed on the dyno and, after running for 10 hours and 54 minutes, was removed, inspected, and prepared for installation on the propeller test stand. That engine first ran on the prop stand, equipped with a three-bladed club propeller, on February 18, 1948.

Final testing and closure

In late 1948, two more major compressor rotor blade failures in different engine assemblies dictated a change to aluminum blades in rows one through twelve (originally magnesium). All of the magnesium vanes in the stator were changed to steel vanes. By the time the contract had come to an end, on June 30, 1949, four engines had been built and run for a total of 610 hours and 43 minutes; and Engine No. 2 had completed a preliminary 50-hour flight rating test in July 1948, while Engine No. 1 was being prepared for an official 50-hour flight clearance test.

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The size of the intake and exhaust are small compared to turbine engines today.

The engines were 90" long, 41" high and 40" wide - not including the aft-pointing exhaust ducts. The dry weight of the engines, including all accessories, was 1,250 pounds, which included the 215 pound regenerator.

By way of comparison, typical World War Two air-cooled reciprocating engines of about 1800 hp were achieving about 0.45 lbs per horsepower per hour at 20,000 feet altitude with liquid cooled engines - slightly better, but at higher installed weights than the XT36.

Still, the engines had not met the Navy's performance expectations; they only developed 803 horsepower, with the regenerator installed, and the specific fuel consumption was 0.664 pounds of fuel per horsepower per hour. (With the dummy regenerator the engine developed 1,039 horsepower but the SFC was 0.770 pounds per horsepower per hour).

These figures were a far cry from the original performance estimates of 1,030 horsepower and 0.510 pounds per horsepower per hour at sea level.

On the day the contract ended, the upgraded A-86B design had been finalized, parts were on order, and Chrysler was confident that that version of the engine would be ready to run by the end of 1949.

One can only speculate as to why the Navy did not extend the contract for another six months at that stage of development, especially since the A-86B version showed promise of meeting or exceeding all of the specifications. One reason may have been postwar military budget cuts; and the emphasis in both the Navy and the Air Force was now on fighter aircraft with more thrust and faster speeds, and fuel economy was of little concern in those early years of the jet age.

In addition, the power output might reach 2,000 horsepower, which was well under what they really were looking for in gas turbine engines. Compare that with a large reciprocating aircraft engine at that time, a well-known device, cheap and easy to make, people know how to do it - they could make 2,220 horsepower with no problems, at a low price. So they just said "Well, we're undersized and over cost."

It was a learning exercise for Chrysler and for the Navy. Fuel economy was better than a typical turbine, but still higher than a reciprocating engine. Brake-specific fuel consumption was maybe double what modern gas turbine engines are doing, because today we're running much higher compression ratios and higher temperatures. The turbines are much more efficient. All of the other accessory drives consume less power.

The aftermath

Chrysler Corporation had benefited immensely from its exposure to gas turbine technology. This was the engine that the guys who ultimately did the gas turbine car cut their teeth on; it was like a gas turbine college, a key step to getting to the gas turbine. They learned things about regenerators or recuperators. They concluded that as the technology existed at that time, not today but at that time, they couldn't get the compression ratios with an axial flow compressor; and that a centrifugal compressor packaged better in a ground vehicle. They cut their teeth and had a chance, at the Navy's expense, to do dozens of burner experiments. In my opinion, it cut four years off the gas turbine project for cars.

Shortly after the XT36-D2 contract came to an end George Huebner's Research Department embarked on a company-funded program to design and build a gas turbine-powered automobile that generated a tremendous publicity for Chrysler Corporation and its products.

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Five years later, in March 1954, Chrysler's first turbine car was introduced to the press. Developing the A-86 / XT36 had laid the groundwork for turbine cars, each with a power plant that was an advancement over the previous design, culminating in the widely publicized program in which 50 Ghia-bodied turbine cars were placed into the hands of 200 randomly selected individuals.

There is an axial flow electric fan in the HVAC system the B-bodies in the 1960s,; that was a spinoff developed by Al Bell, one of the engineers on the gas turbine car. He was looking and suggesting ways of making smaller and cheaper. It had manufacturing problems, because of the tip clearance, but it was there.

There were a lot of very dedicated and smart guys that worked on this and honed their skills; I don't think that there's been enough credit given to them, and the growth of knowledge and the product-oriented research and techniques that they brought along. That's the story to me.

I always thought that the place in terms of ground transportation for gas turbine was over the bus, a Greyhound bus, because they run at high speeds. They run continuously. They get regular, careful maintenance, and they don't have long periods in which they idle. I always thought we somehow or other missed it by not trying to do one for an over the road bus.

Recuperators or regenerators, in aviation

In the long history of aircraft gas turbine engines, only two other companies have ever built or tested engines with an integral recuperator. In England, Bristol built the 2,000 horsepower Theseus turboprop in 1944, but abandoned the program two years later due to complexity and weight problems. In this country, Allison, a division of General Motors (at that time), designed a 4,000 horsepower turboprop in 1964; after one engine was built and tested, the contract was cancelled. Allison also adapted a recuperator to their 250 horsepower Model 250 turbo-shaft engine in the mid-1960s but was not further developed.

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As a result of improved aerodynamics, advances in metallurgy, and overall design improvements, turbine engineers have been able to lower SFC values without resorting to heat exchangers. A typical modern turboprop, the 1,020 horsepower Pratt & Whitney Canada PT-6, has an SFC of 0.62. This is slightly better than the XT36, which required a heat exchanger to reach an SFC of 0.66. Furthermore, the PT-6 is 19 inches in diameter, 76 inches long and weighs 490 pounds; whereas Chrysler's XT36 was 40 inches in diameter, 90 inches long and weighed 1,250 pounds.

Sadly, to the best of the author's knowledge, none of the Chrysler XT36-D2 turboprop engines exist today nor have any of its components survived. It remains today only a memory in the minds of the few surviving engineers, designers and technicians who worked on that now little known and almost forgotten power plant.

Epilogue: where did the engineers go?

Chrysler Corporation's Research Department was staffed by a small group of talented and dedicated engineers, metallurgists, physicists, designers, draftsmen, technicians, and mechanics, all of whom made major contributions to a number of challenging problems.

Robert Pauley, part of the Research Department from 1953 on, reached the level of Design Supervisor before joining Williams International in 1981.

George J. Huebner, Jr. was born in Detroit on September 8, 1910. After graduating with an engineering degree from the University of Michigan, he joined Chrysler Corporation in 1931 as a research engineer. Five years later, he was named assistant chief engineer of Chrysler's Plymouth Division, staying at that job until 1939, when he was promoted to chief engineer of the Research Division. There he directed the design and development of the XI-2220 and the XT36-D2 engines.

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In 1952, George Huebner served as executive engineer of Chrysler's Missile Division, while still assigned to the Research Division. In 1955 he became Director of Research and was the driving force behind Chrysler's automotive gas turbine program. Often referred to as "The Father of the Automotive Gas Turbine," Huebner retired from Chrysler in 1976. After retirement he served as President of SAE and was awarded SAE's prestigious L. Ray Buckendale award for his contributions to vehicle engineering. He also received an award from the Society of Mechanical Engineers for "Gas Turbine Leadership."

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Huebner later served as a consultant for the Assistant Secretary of Defense for Research and Development and was also named CEO of the Environmental Research Institute of Michigan. He died in 1996 at the age of 86.

William Irvin Chapman, a graduate engineer from Iowa State University, had joined Chrysler in 1941, and during the war years worked on the XI-2220 aircraft engine. He tested the Japanese Ne-20 engine for the U.S. Navy and later was in charge of the dynamometer and prop-stand testing of the XT36-D2 turboprop. When that program came to an end, Chapman played a key role in the design and development of Chrysler turbine engines. In 1964 he left Chrysler to work for Ford, helping to develop Ford's Model 707 truck and industrial gas turbine engines. He left Ford in 1974 to join Williams Research Corporation where he worked on a variety of missile and aircraft turbine engines.

Chapman retired from Williams in 1993 as manager of the Preliminary Design Department. He maintains an active interest in gas turbine engines after a 60 year career devoted to the subject. The author is deeply indebted to AEHS member Bill Chapman for his assistance in the preparation of this article.

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Dr. Sam Williams
, a graduate of Purdue University, joined Chrysler in 1942 as a research engineer. He worked on the preliminary design of the XT36-D2 and later directed the design of the upgraded version of that engine, the A-86B. He was very smart and very motivated.

When the XT36-D2 program ended in mid-1949, Williams was promoted to manager of the design department where he was responsible for the design of Chrysler's first automotive gas turbine engine, the A-249. As conceived by Dr. Williams, the A-249 featured a rotating disc-shaped regenerator combined with a small gas turbine power plant that was compact enough to fit under the hood of a passenger car.

Sam Williams left Chrysler in 1954 to form the Williams Research Corporation. Under his leadership that company grew rapidly building turbine engines for target drones and cruise missiles. Today, Williams International is a major manufacturer of gas turbine engines including the FJ33 and FJ44 turbofans widely used to power a number of entry level business jets. For his accomplishments in the small gas turbine field Dr. Williams has been awarded the Collier Trophy (1979), the Wright Brothers Memorial Trophy (1988) and has been inducted into the National Aviation Hall of Fame (1998).

John Jones, a graduate of Case Western Reserve University, did the analytical work on the axial-flow supercharger for the XI-2220 that led to the design of the XT36-D2 engine. He left Chrysler in 1954 and, along with two other Chrysler employees, joined forces with Dr. Sam Williams when the Williams Research Corporation was formed. He retired from Williams International as Vice-president of Engineering and died on July 9, 2000 at the age of 81.

This page was based on a presentation by Robert Pauley and Al Bosley, and a phone interview with Al Bosley, who said: "Robert Pauley is the guy that saved the material and did the research. I was only an editor, colorist, and slide maker. He did the heavy lifting and I did the fun work."

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