The Early History of the Axial Type of Gas Turbine Engine

41 1

The Early History of the Axial Type of Gas Turbine Engine

By Hayne Constant, M.A.

Introduction. In this lecture, I have attempted to put on record a history of ideas. I have tried to show how the outlook of those responsible for directing the early researches in this country on the axial type of gas turbine engine developed as the work pro- ceeded and how the progress achieved reacted back on their state of mind. It has been my object to emphasize the reasoning behind each technical step that was taken and then to show how subsequent developments proved or disproved the validity of this reasoning.

Although this treatment of the history of a very interesting technical development is by no means intended as a post- mortem, it is my hope that it may assist engineers in avoiding in their own work the mistakes that have been made in the past.

EarIy Histmy. The first practical proposal to use a gas turbine as an aeroplane power plant was made by the Royal Air- craft Establishment in July 1926, when Dr. A. A. Griffith outlined his aerofoil theory of turbine design.

In October of that year, a conference was held at the R.A.E. at which Dr. Griffith put his proposals to a small committee from the Air Ministry and the Aeronautical Research Com- mittee. This conference expressed itself unanimously in favour of putting in hand preliminary experiments to verify the theory. Accordingly, a test rig consisting of a single-stage turbine driving a single-stage axial compressor-both axial flow-was designed. The rotor was operated by sucking air through it and measuring the total head losses: this was the first occasion on which free vortex flow blades were used. The unit was tested in 1929 and a stage efficiency of 91 per cent was achieved.

At the same time, the first wind tunnel tests on cascades of compressor and turbine blades were made at R.A.E. These were completed in 1927, but the results* were not available in time to be incorporated in the turbo-compressor unit.

In spite of the success of this experiment, approval for the construction of a turbine could not at that time be obtained and no further work was done until several years later-in 1936. During those seven years no furrher progress in the arts relating to the gas turbine had been made, so what was believed possible in 1936 could equally well have been done earlier. All that had changed was mans outlook.

In his aerofoil theory of turbine design, Dr. Grf i th had established the basic principles of the design of aerodynamic compressors and turbines of the multistage axial or radial flow type. His theory, in fact, could be applied to any rotary mechanisms in which the working fluid was deflected by blades of aerofoil shape.

Several important conclusions resulted from his work. It was clear, for example, that it should be possible in a compressor or turbine to attain small stage efficiencies of at least 90 per cent. This conclusion was borne out by the results of the rig tests already described. Again, it appeared that it was desirable, both in the interests of efficiency and of control of the working fluid, to arrange that the air flowed in a free vortex in the annular space between one blade row and the next. His study of the range of incidence over which a cascade of aerofoils could operate led to the conclusion that in a multistage compressor of high pressure ratio there would .be inefficient operation and danger of stalling when operating under conditions considerably different from those for which the blading was designed.

This led to the conception of the compound turbine engine in * HARRIS, R. G. and FAIRTHORNE, R. A. 1928-9 Technical Report

of the Aeronautical Research Committee, No. 33, vol. 1, p. 286, Reports and Memoranda, No. 1206.

The Technical Background in 1936.

which compression was carried out in a number of mechanically independent stages, each stage being driven by its own turbine. In such an arrangement, the rotational speeds of each stage adjust themselves to the running conditions in such a way as to reduce the range of incidence over which the blades have to operate and thus decrease the danger of stalling. Finally, it was concluded that to obtain satisfactory operation under part load conditions, it was desirable, but not essential, to use a power turbine in paralleI with the compressor turbine.

There was a small amount of earlier work on which it was possible to draw. Sir Charles Parsons had designed and operated a number of axial compressors from 1904 onwards, but, as these had stalled blading, their efficiency was not high enough for use in a gas turbine.

A considerable amount of experience had been gained at the R.A.E. on exhaust gas turbo-compressors. This work had shown that there was little difficulty in operating turbine wheels running at high temperatures and high tip speeds. For example, considerable bench and flight testing was done on single-stage impulse turbines with gas temperatures of 900 deg. C. (1,652 deg. F.) and tip speeds of over 1,000 ft. per sec. and the only serious trouble that remained to be overcome was overheating of the turbine bearing. The efficiency of these turbines was, however, only about 60 per cent-due partly to their design not being based on aerodynamic knowledge and partly to the limitations in blade design imposed by their method of construction.

In centrifugal compressors, efficiencies of 75 per cent had been attained at pressure ratios of 2/1 but at higher pressure ratios the efficiencies were much lower. Materials were available having reasonably good high-temperature properties. Thus in Hadfields ERA/ATV we had a material which had a maximum stress of about 30 tons per sq. in. and a useful creep strength of 5 tons per sq. in. at 700 deg. C. (1,292 deg. F.). It was with this background that in 1936 we set to work on the gas turbine engine at the R.A.E.

At that date, we believed that the axial type of compressor was inherently capable of higher efficiencies than the centrifugal type. But practically no knowledge of the capabilities and limitations of axial compressors was available. On the other hand, we felt confident from our experience of exhaust gas turbines that no insuperable difficulties would be encountered in the turbine end of the machine. Accordingly, we decided that the immediate objective to be attained was a satisfactory axial compressor.

The first step was to produce an experimental multistage axial compressor to test the theory that had been built up and to see to what extent the difficulties that had been anticipated were real. Accordingly, we designed a small eight-stage axial compressor which later became known as Anne.

The compressor as first constructed is shown in Fig. 1. The aerodynamic design was based on an assumed adiabatic stage efficiency of 90 per cent.

The design conditions were :-

Anne.

Tip speed . . 750 ft. per sec. Mean axial velocity . 500 ft. per sec. Mass flow . Tip diameter . . 6.0 inches Rotational speed . 28,600 r.p.m.

. 3 lb. per sec. at N.T.P. entry conditions

The blades were designed for 50 per cent reaction at the inner radius with a twist to give free vortex flow. Blade camber was 45 deg. at all radii both on rotors and stators. The blade profiles design was that known as R.A.F. 27 thickened 10 per cent

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412 DEVELOPMENT O F THE INTERNAL COMBUSTION TURBINE

on a circular arc backbone. The performance of this blading was deduced from the wind tunnel tests described in R. and M. 1206. The mean pitch/chord ratio was 1-30 and the average axial clearance between the blade rows 11 per cent of the chord.

To prevent stalling when starting and when running under conditions greatly different from the design conditions, each stage had a large number of bleed holes through which a controlled amount of air could be blown off from each stage. Bleed holes were also provided on the outlet diffuser.

Owing to the small size of the blades it was anticipated that considerable difficulty would be experienced in devising a satisfactory method of attaching them to disk rims running at 500 ft. per sec. I t was accordingly decided to machine the blades integral with the disks. This difficulty was imaginary and was due to our having inadequate experience of the mechanical design of machinery of this kind. It had a serious influence on the aerodynamic design, for in order to get the profiling tool between the blades it became necessary to pitch the blades much farther apart than was desirable. This mistake of compromising the aerodynamic design because of real or imaginary mechanical limitations was repeated again and again during the following years. It is a lesson which will only be learned by people with as much confidence in mechanical as in aerodynamic design.

When Anne was first tested, the only motor available could be operated only from half to full speed. It was, therefore, impossible to motor the compressor slowly while its mechanical operation was corrected. The result was that immediately after its first start, an oil seal rubbed and caused one of the disk wheels to overheat. The blades carried by this wheel rubbed on the outer casing and broke off and the remaining blade rows were stripped. Thus over eighteen months work was lostin 30 seconds.

The causes of this minor tragedy were the imperfect mechanical design of the compressor and the lack of suitable testing equipment. A great deal of research, even in aerodynamics, requires the services of competent mechanical engineers and these can only do their job effectively if they are kept abreast of development by actually engaging in design, manufacture,

and testing. The lack of testing equipment in 1936 reflected the general state of neglect into which our Service research establish- ments decay when the stimulus of a major war is withdrawn.

After this accident, immediate steps were taken to redesign the compressor. During the interval since the f is t design had crystallized, certain rumours, information, and changes in outlook had occurred. We had heard of the success of the Brown Boveri axial compressor, which could be started and operated satisfactorily without air bleeds. As we had been very nervous that the air bleed holes in Anne would have a spoiling effect on the air flow when not in operation and might in addition cause mechanical failure by allowing bits of swarf to pass through the blading, we decided, in the light of the Swiss information, to eliminate all air bleeds in the new design. Later experience proved this to be a retrograde step and air bleeds were reintroduced in other compressors several years afterwards.

Owing to the difficulties that had been experienced in cutting bladed wheels from the solid, we decided to make the blades separately and mount them on the wheel rims on circular bases so that the stagger angle was adjustable.

We had also heard from Switzerland that to get maximum efficiency it was desirable to operate with an axial clearance between the blade rows of not less than one-third chord. We were at the same time concerned about the danger of the blade wakes from one row inducing vibrations in the succeeding row of blades if they were placed too closely together. We, therefore, decided to reduce the blade chord from 0.55 to 0.438 inch, thus increasing the axial clearance to a mean value of 37 per cent of the chord. Later experience has shown that this was a mistake and that a higher efficiency could be obtained by operating with smaller clearances.

The result of this change in the blading had a most serious effect on the aerodynamic performance, since the resulting increase in pitch/chord ratio from a mean value of 1.3 to 1.63 reduced the predicted pressure ratio of the unit from 4/1 to just under 3/1. The layout of the redesigned compressor is shown in Fig. 2.


EARLY HISTORY O F THE AXIAL TYPE

1.0-

e

21.5- 5

z 2

3 R

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2 I . O

0.5

0-

413

lO.oOO\

l8.000-\

l6.000-, \

14,000

12.m-

0 0.5 I * 0. 1.5 2.0

Fig. 2. Axial Compressor Anne as Tested

t P

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- I C A M B E K ~ .

I ANGLE I I

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Fig. 3. Nomenclature for Compressor Blading

The aerodynamic design was again based on free vortex blading with 50 per cent reaction at the h e r radius. The blade profile was R.A.F. 27 on a circular arc camber, with a thickness/chord ratio of 13 per cent. The definitions of camber and entry angle, etc., are indicated in Fig. 3, showing the general nomenclature used for axial compressors.

Anne ran successfully for the first time towards the end of 1938. The technique of testing had been greatly improved and the only mechanical troubles experienced were in the t h r u s t bearing. This bearing had to take a thrust of about 250 lb. at 27,000 r.p.m. As is usually the case with a new ball bearing installation, considerable trial and error modifications had to be made before satisfactory operation was obtained. We learnt from these experiences the necessity in a new ball bearing installation of keeping a constant watch on the temperature of the stationary race. Many failures have been prevented by stopping units immediately an unexplained rise in temperature occurs.

Blade tip clearance indicators were used for the first time on t h i s compressor and proved very successful. The indicator con- sists simply of an insulated needle which is screwed in through the casing until it makes contact with the tips of the rotating blades and thus completes an electrical Circuit.

T h e characteristics of h e are shown in Fig. 4. The Der- Fig. 4. Characteristics of h e

,

~~

formance was not quite up to predictions but Gas sufliciehtly Values plotted are for standard entry conditions: 14 lb. per sq. in. a t promising to encourage us to push ahead with the development 288 deg. C. (550.4 deg. F.)


414 DEVELOPMENT O F THE INTERNAL COMBUSTION TURBINE

Fig. 5. Axial Com

Ruth. The information and ideas arising out of the design and testing of Anne were applied in 1939 to a new compressor called Ruth whose mechanical design and manufacture were due to the Fraser and Chalmers works of the General Electric Company.

The principal step forward in this compressor was the reintroduction of a small pitch/chord ratio. I t will be recalled that this ratio in Anne was undesirably large owing to the blade chords being reduced to give greater axial spacing between the blade rows. On Ruth, the mean pitch/chord ratio was 1.03 compared with 1.63 on Anne. The result of this change was a considerable increase in the pressure rise per stage, so that in the six

Annes eight stages. Fig. 5 shows a general arrangement of the compressor. The

six rotor stages were mounted on a drum built up from rings pulled together by a central shaft. The reasons for changing from the disk to the drum type of construction will be referred to later when dealing with complete turbine engines.

2.5

2.

3 i**

T h e design conditions for Ruth were :- Maximum tip speed . 650 ft. per sec. Mean axial velocity . 430 ft. per sec. Mass flow . . 6.0 Ib. per sec. at N.T.P. entry

conditions Tip diameter . . 8.0 inches Rotational speed . 18,700 r.p.m.

The basic aerodynamic design was similar to that of h i e . The performance of the compressor is shown in Fig. 6. The

chief points of interest are the higher pressure ratio per stage and the serious falling off in performance at speeds above the design speed. This deterioration was expected and was, I believe rightly, attributed to compressibility effects in the low- pressure stages of the compressor. At 19,000 r.p.m. the maximum Mach number* was 07 and the highly cambered blades used

* The Mach number or coefficient is the ratio of the actual to the acoustic velocity in the fluid.

,. ,. ,.

Fig. 6. Characteristics of Ruth Values plotted are for standard entry conditions : 14.7 lb. per sq. in. at 288 deg. C . (550.4 deg. F.) abs. Total head efficiencies are about 8 per cent higher than the static head efficiencies shown.


EARLY HISTORY OF T H E AXIAL TYPE 415 would not operate much above this figure without a compressibility stall occurring.

A number of other axial compressors were constructed which gave us detailed information on a number of points but as they did not lead to radical changes in our outlook they will not be referred to here.

In 1937, following a recommenda- tion by the Engine Sub-committee of the Aeronautical Research Committee under the chairmanship of Sir Henry Tizard, the Air Ministry authorized the Royal Aircraft Establishment to start work on the problem of the gas turbine. As the R.A.E. was not equipped to carry out large-scale manufacture, it was arranged that detail design and manufacture should be carried out for the Establishment by the Metropolitan-Vickers Electrical Company.

The first scheme considered (designated A) involved the use of centrifugal compressors. These, however, were quickly abandoned and all further work was concentrated on the axial compressor type, because it was believed that this type could give higher efficiency with a lower frontal area and less bulk. The first requirement was for a plant to develop brake power on the test bed for a demonstration of the practicability of the gas turbine.

We were not aiming at a power plant in which weight and bulk were reduced to the limit, for this would have involved us in prolonged development work which neither the R.A.E. nor Metropolitan-Vickers could at that time have undertaken. We were, however, anxious to produce a layout which was inherently of a type on which a compact and light power plant could later be based. It was for this reason that we sacrificed compressor and turbine efficiency in order to use the smallest possible number of stages and ran our compressors at higher Mach numbers than the attainment of maximum efficiency would require.

Early Turbine Schemes.

I have already referred to the difficulties which we feared would be encountered if we tried to develop too high a pressure ratio in a single compressor.

We, therefore, started off with the idea of a compound engine with two mechanically independent compressors. A large number of alternative arrangements of components had to be investigated. One of the earliest of these is shown in Fig. 7. This was a double compound engine with a power turbine in parallel with the compressor turbine. These early studies brought us face to face with the conflicting requirements of mechanical and aerodynamic simplicity. The conflict exists in every turbine layout we have considered and the ideal will not have been reached until we can devise an arrangement in which both sets of requirements are identical.

In the layout shown, the aerodynamic requirements have been given priority, the air being subjected to (as we then believed) as few unnecessary bends as possible, and the frontal area reduced to a minimum. But the mechanical complication of two concentric shafts was more than we could face, and the scheme was abandoned at an early stage. We decided to avoid the mechanical difficulties of the concentric arrangement by dispersing our units as shown in Fig. 8. In this layout, the aerodynamic requirements were sacrificed in order to givc a simple mechanical arrangement in which each compressor was directly driven by its own turbine so as to form a simple independent unit. The price that was paid for this simplicity was a flow path for the working fluid which involved no less than twelve right-angled bends.

In those days, we were supremely confident of our powers to control the working fluid but not so sure in our knowledge of mechanisms. The consequence was that we tried to evade the mechanical problems but gave ourselves some serious aerodynamic difficulties as a result. This outlook persisted for some time.

It was finally decided not to fix the final arrangement of com-

Fig. 7. Arrangement of Coaxial Compound Turbine Engine


COMPRESSOR

POWER TURBINE

Fig. 8. Arrangement of Dispersed Compound Turbine Engine Pressure ratio: 5/1.

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EARLY HISTORY O F THE AXIAL TYPE 417 ponents until tests had been made on a single turbo-compressor unit. The object of these tests was to obtain confirmation that the mechanical details of design were sound and that the actual performance obtained confirmed that predicted from a knowledge of the performance of the component parts. This decision, which was mainly due to lack of confidence in our own judge- ment, turned out to be a ven lucky one since it gave us time to appreciate the aerodynamic defects of the complete scheme. A unit known as the B.10 was therefore constructed.

The B.10 Turbo-compressor (Betty). The B.10 turbo- compressor represented the high-pressure unit of a complete compound engine rather similar to that already shown in Fig. 8. It consisted of a 9-stage axial compressor driven by a 4-stage turbine. The layout is shown in Fig. 9, but the combustion chamber connecting compressor delivery to turbine intake has been omitted in the diagram. The unit was designed to be self- running, but delivering no useful power. The compressor was first tested separately by driving it with a steam turbine. The testing was carried out by Metropolitan-Vickers.

This compressor had blading based on the same general principles as our earlier compressors, but with refinements em- bodying recent increases in our knowledge of flow past blade cascades.

The mean value of pitch/chord ratio was 0.80. The tip speed at entry was 566 ft. per sec. and at exit 460 ft. per sec.; the mean axial velocity 356 ft. per sec.; the rotational speed, 7,000 r.p.m.; and the designed mass flow 16-4 lb. per sec. at N.T.P. entry conditions. The compressor was tested in 1939 and gave a very good performance. The test results are shown in Fig. 10.

lb !II MASS FLOW-LB. PER SEC.

Fig. 10. Characteristics of Betty Values plotted are for standard entry conditions: 14.7 lb. per sq. in.

at 288 deg. C. (550.4 deg. F.) abs.


418 DEVELOPMENT OF THE INTERNAL COMBUSTION TURBINE This was the first axial compressor to be tested with a

reasonably high Reynolds number and the satisfactory aerodynamic and mechanical results obtained gave us confidence that we were working along the right lines.

After the turbine had been tested separately with steam and the combustion chamber had been developed, the complete unit was assembled and tested in October 1940. The only troubles experienced during its operation were the bearing failures which we have now come to recognize as being nearly always as- sociated with a new bearing installation.

There were a number of mechanical features in the B.10 unit to which it is worth while drawing attention. The first of these is the drum type of rotor used in both the turbine and compressor. The principle reason for the use of a drum rather than a disk type of construction was to obtain uniform expansion and contraction of the rotor and casing. We feared that a disk with its higher heat capacity would cool down less rapidly than the casing and cause blade fouling when the unit was stopped. Although no trouble due to this cause had occurred on the compressor Anne, the test conditions had been far less onerous than those occurring in the B.lO. It has yet to be proved that these fears were unnecessary, but experience is gradually accumu- lating which strongly suggests that no troubles of this sort will occur.

We were very reassured to find that satisfactory operation with freedom from distortion could be obtained with a red-hot rotor -inlet temperature 675 deg. C. (1,247 deg. F.)-running at high speed. Although we had had little trouble with exhaust gas turbines operating at considerably higher speeds and temperatures, the clearances allowed were much greater than could be tolerated in the high efficiency reaction blading of the B.lO.

Water-cooled bearings were used with success in this unit, but were abandoned in later machines in favour of the simpler air cooling.

The principle aerodynamic lesson that we drew from the tests was that the losses occurring in collecting elbows and volutes were more than could be tolerated. This confirmed the results obtained from some volute tests which were carried out while the B.10 was under construction. It became quite clear that for aircraft applications, where space was limited, our decision to avoid mechanical complication by the introduction of features which were aerodynamically undesirable, was unsound. In a gas turbine, whose performance depends so intimately on the various losses suffered by the working fluid in its passage through the machine, there must be no compromise with the aerodynamic requirements.

The appreciation of this point completely changed our outlook on design and we abandoned our earlier conception of a dispersed double-compound engine.

We had then to decide on an alternative arrangement in which a smoother path was provided for the working fluid. It was clear to us that the layout must be such that all the machinery was coaxial, so that no collector elbows or volutes would be required. The principal point at issue was whether compression should be carried out in a single compressor or whether we should need to compound and use two mechanically independent coaxial compressors in order to get sufliciently flexible operation to obtain easy starting.

The decision we were required to make was a very difficult one. The pressure ratio for which we were designing was only 5/1 and considerable evidence had accumulated that up to this ratio it should be possible to start comparatively easily-without stalling the compressor-without resorting to the complication of compounding. On the other hand, this unit was an experimental prototype which we hoped would show the way to further developments along similar lines. And since these developments would naturally be directed towards the use of higher pressure ratios, we were reluctant to build into the engine a feature which might prevent its development to those higher ratios.

The decision that was reached was again to avoid mechanical complication. We decided to do all our compression in a single compressor and to postpone to the future the problems of compounding. I have regretted this decision ever since.

This refusal to face mechanical problems is all the more serious when it is remembered that this development was the responsibility cf the Engine Department, which certainly had more

mechanical than aerodynamic knowledge. It may be that our knowledge of mechanisms was sufficient to make us aware of the difficulties that had to be overcome, while our comparative ignorance of aerodynamics allowed us to accept problems in this field with equanimity. We, therefore, shirked the difficulty we could foresee and plunged lightheartedly into the aerodynamic morass from which more experienced aerodynamicists might have recoiled.

It is interesting to speculate on the form of gas turbine that would have been devised by a band of aerodynamicists without mechanical experience. Would it have been a weird contraption of cogs and pulleys with everything arranged for the comfort and guidance of the all-important working fluid?

I recall very well a period in 1938 when the difficulties likely to be encountered by differential thermal expansion and distortion overshadowed my thoughts like a nightmare.

It will, I think, be agreed that there was some justification for these forebodings, for an engine of the size we were con- templating would increase in length by about 1 inch when hot. Again, the blade clearances at which we wished to operate would have been completely taken up by a change of only 50 deg. C. (90 deg. F.) in the local temperature of certain parts of the engine. Further, the degree to which the materials used could flow plastically and thus relieve thermal stresses while avoiding permanent distortion was not known.

The Nightmare of Thermal Distortion.

WATER

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EARLY HISTORY OF T H E AXIAL T Y P E 419

For these reasons, very considerable thought was given to the problem of reducing relative expansion as much as possible. Many schemes-some of terrifying complexity-were considered, and as an example, the layout shown in Fig. 11 may be of interest. In this arrangement it was proposed to insulate the main structure of the turbine from the hot gases by using a water-cooled internal liner. The rotor was to be similarly insulated by water jackets between the disk rims, the water being fed in through the stator blades.

An alternative method of preventing heat flow into a turbine disk or drum depends on passing the cooling medium along axial slots beneath the blade roots. A rig was made up to measure the effectiveness of such an arrangement, using air as the cooling medium, and it was found so satisfactory that it was later incorporated in the F.2 engine. The idea is quite straight- forward and simple and is illustrated in Fig. 12.

Another example of detail design to reduce thermal stresses is the double cone type of drum end piece used with success on the B.10 and on later turbines. The design of these end pieces caused a great deal of worry, since the calculated elastic stresses were over 100 tons per sq. in.; but the plastic yield saved the situation.

Fig. '12. Detail of Air-Cooled Disk Rim

The D.11 Gas Turbine (Doris). The new unit was known as the D.ll gas turbine and a general arrangement is shown in Fig. 13. The unit consisted of a 17-stage compressor directly driven by an 8-stage turbine and power was obtained from a 5-stage low-pressure turbine taking gas from the exhaust of the high-pressure turbine.

In order to fit the D.ll into the background of what we then believed and now know to be possible, the generalized performance curves for propeller turbines under ground level static conditions are shown in Fig. 14, with the approximate design point for the D.11 indicated.

These curves are based on the following assumptions :- Compressor small stage efficiency . . 87 per cent Main turbine efficiency . - 87 9, Power turbine and exhaust efficiency . 88 ,, Combustion efficiency . - 98 3, Combustion pressure loss . . 2 lb. per sq. in. Jet velocity . . 500 ft. per sec.

Total head efficiencies are given. It will be seen that we were attempting to do something quite

modest compared with the tremendous possibilities that lie before us towards the bottom right-hand corner of Fig. 14.


420 DEVELOPMENT OF T H E I N T E R N A L C O M B U S T I O N T U R B I N E

0.3 ! I 1 I I 0 40 80 I20 160 ZM) 240

SPECIFIC EQUIVALENT B.H.P.-B.H.P. PER LB. OF AIR PER SEC Fig. 14. Generalized Performance Curves for Propeller

Engines Propeller turbine : sea-level static. T3[ = Total head temperature at combustion chamber outlet. P3[ = Total head pressure at combustion chamber outlet. Plt = Total head pressure at compressor inlet.

The compressor was first constructed and tested in 1941. There is no need to go into the details of the blading since the tests showed this to have undesirable features. Fig. 15, Plate 1, shows the compressor opened up.

The test results are shown in Fig. 16. Although owing to its many stages this compressor gave a considerably higher pressure ratio than we had ever obtained before, its charaaer- istics at high speed were unsatisfactory. The sudden jump from a compressor of nine stages to one of seventeen was an ambitious advance and the poor results obtained showed that we had out- stepped our knowledge. This was the first occasion on which it had not paid to be too bold, for our previous failures had been due to lack of courage.

The trouble on Doris was to some extent due to the Mach number at entry being too high but more important was the fact that we assumed in the design too large a thickening of the rotor and casing boundary layers as the air passed through the compressor.

The result of this was that the effective flow path at the high- pressure end of the compressor was greater than we had designed for. As a consequence of this the mean axial velocity was reduced and the blade incidence increased to such an extent that blade stall occurred at a comparatively high mass-flow. The compressor surge line was thus swung over in the direction of higher mass-flows and operation on those parts of the characteristic which normally gave the highest efficiency, became impossible.

The results of the compressor tests showed that reblading of a number of stages would be necessary before the mating of compressor and turbine would be sufficiently good for self- running of the complete unit to be obtained.

At this time a jet propulsion project, which had been proceeding in parallel with the work already described, reached a stage at which it was decided to give it priority over the D.ll. The reblading of the compressor was accordingly abandoned for the time being and little work was done on the rest of the unit.

I I I I I

Fig. 16. Characteristics of Doris Values plotted are for standard entry conditions : 14.7 Ib. per sq. in.

at 288 deg. C. (550.4 deg. F.) abs.

Although I have made incidental references to the complete turbine engine, t h i s history has so far consisted mainly of the story of the axial compressor. To preserve continuity, I will continue this story and then return to a fuller consideration of the complete engine.

The next compressor was designed for the jet propulsion project to which I have just referred and to which I will return later. The design conditions for this compressor, known as Freda, were as follows :-

Mass flow . . 50 lb. per sec. at N.T.P. entry conditions

Pressure ratio . . 4/1 Rotational speed . 7,390 r.p.m. Number of stages . 9 Maximum tip speed . 718 ft. per sec. Mean axial velocity . 500 ft. per sec. Tip diameter . . 22.2 inches

Fredu.

The blading was of free vortex design using as usual R.A.F. 27 profiles on a circular arc backbone, with a mean thickness/chord ratio of 13 per cent. The mean pitch/chord ratio was 0.90 at the outer and 0.68 at the inner radius. The rotor blades were all similar and set at the same angle, the progressive reduction in height towards the higher pressure end being achieved by cutting the blade tip to the appropriate length. The stator blades were all similar to each other, length adjustment being made by cutting off the tips; they were untwisted.


EARLY HISTORY OF THE AXIAL TYPE 42 1 The compressor was first tested separately and gave an

excellent performance. The test results are shown in Fig. 17. The performance of Freda was the best so far obtained from

an axial compressor. Its good performance was due to a number of factors. In the first place it operated at a higher Reynolds number than any previous compressor. Its blade aspect ratio was also higher than any earlier compressor except Doris; and Doris, as we have already seen, suffered from a number of ail- ments which masked any benefit it might have derived from this design feature. Freda also had blades produced by a new press- ing process which gave both a good finish and very accurate and consistent profiles.

Fig. 17. Characteristics of Freda Values plotted are for standard entry conditions: 14.7 lb. per sq. in.

at 288 deg. C. (550-4 deg. F.) abs.

It is a characteristic of axial compressors that the cumulative effects of small deviations from design in the low-pressure stages may have a serious effect on the performance of the high- pressure stages, and this will react on the performance of the whole compressor. This trouble is, of course, most serious in high-pressure compressors. In Freda our earlier experience, and a little luck, resulted in our estimates of the progressive change in the condition of the boundary layers through the compressor being more accurate than before. The result of this was that the high-pressure blades were in fact subjected to conditions very nearly in accordance with those for which they had been designed.

Sarah. The next step in the axial compressor development was an attempt to improve the performance of the best existing

axial compressor (Freda) by adding further stages to raise its pressure ratio. The new compressor (Sarah) was manufactured by Armstrong Siddeleys as part of a jet propulsion turbine engine known as the A.S.X. The blading of Sarah was in two parts : the high-pressure part was identical with that of Freda; the low-pressure part consisted of five stages of blading generally similar to the Freda blades but designed to have constant reaction at all radii instead of reaction increasing with radius, as is implied in the free vortex blades used on all our earlier compressors. This first departure from free vortex blading deserves some

comment. We had for some time suspected that the losses in a forced vortex might not be appreciably different from those in a free vortex in which the angular momentum was inversely proportional to the radius. If this proved to be true, it might under certain conditions be preferable to use a forced vortex. For example, with constant reaction at all radii the work input at the blade roots can be increased so that a higher pressure rise per stage can be achieved. Since, however, for a given tip Mach number the Mach number at the root is greater on constant reaction blades than on free vortex blades, a smaller thick- nesslchord ratio, i.e. a larger chord, has to be used on the constant reaction blades. Although this increases the weight and tends to reduce the advantage of the higher stage-pressure rise, there are occasions on which the balance of advantages may be in favour of such blading. The tests on Sarah were not con- clusive, but there were no grounds for deducing from them that constant reaction blading had a lower efficiency than the free vortex type.

The general conclusion that was drawn from these results and from the various other researches that were proceeding on the same subject, was that it was possible to impose on the flow through a compressor or turbine a forced vortex having any angular momentum distribution over a comparatively wide range without serious changes in the blading efficiency. The most suitable distribution to use would depend on the design conditions.

The leading particulars of Sarah were as follows :- Maximum tip speed . 714 ft. per sec. Mean axial velocity . 490 ft. per sec. Mass flow . . 50 lb. per sec. at N.T.P. entry con-

ditions Tip diameter . 20.5 inches Rotational speed . 8,000 r.p.m. Mean pitchlchord ratio 1.24

The compressor tests gave very good results and these are shown in Fig. 18. The high aspect-ratio blade and low Mach numbers are the principal reasons for its good performance at high-pressure ratios.

Blade Stresses. At this point it is convenient to refer to a number of special problems that were always with US. The first of these is the question of permissible blade stresses.

The development of the gas turbine has been very much hampered by lack of knowledge of permissible blade stresses. Both the compressor and turbine blades are subjected to a tensile stress and a steady gas bending stress on which is super- posed a fluctuating bending stress due to the effect on the air flow of stationary members in front of or behind the blades. The tensile and steady bending stresses can be calculated with reasonable accuracy but we were-and to a large extent still are -in great ignorance concerning the magnitude of the forcing impulses set up by the adjacent rows of fixed blades, entry spider arms, etc.

The problem of studying the effect on the life of a blade of changes in the mean stress at which it is run is complicated by the fact that a blade has a considerable number of modes of vibration, each of which involves a resonance speed. The mean stress that a blade can stand is therefore affected by the nearness to resonance at which it is run.

This problem of blade stresses represents probably the biggest gap in our knowledge of turbine engines. It is not a gap which will be filled by general development experience since one might run an engine for years without trouble and then wreck it by a few hours operation right on one of the blade critical speeds.


422 DEVELOPMENT OF THE INTERNAL COMBUSTION TURBINE

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Speed, Bending Centrifugal Material r.p.m. stress, stress,

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1

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Freda . 7,390 2.70 1.43 Al Betty Doris * . I 7,300 I i::: I !;: 1 Sgel 7,000 E.5 . I 171100 I 2.15 I 2.13 I A1 Ruth . I 24;OOO 1 3.97 1 2.54 Al Alice . 8.000 1.66 0.33 ~ ~~ . Anne 24;ooo 0.93 1.45 Al Sarah : 1 8,OOO 1 2-3 I 1.6 I A1

The Combustion Problem. Combustion work was started in a very crude way in 1936. My first fuel pump was a six-cylinder Bosch Diesel engine pump feeding a common gallery, and this supplied Diesel oil to a single orifice.

We knew that we could get a shorter combustion chamber by using vapour instead of liquid injection but we anticipated that serious difficulties would be encountered in attempting to steer a middle course between the Scylla of cracking and the Charybdis of priming. These fears were later confirmed by the troubles met with in the Power Jets vapour injection schemes. Having decided not to use vapour injection we were left with the alternatives of either directing the fuel upstream with solid or atomized injection, or downstream with a swirl or other form of atomization, or some combination of the two systems.

After short trials of both systems, we decided to concentrate on upstream solid injection since we found that, with the par- ticular arrangement of atomizing jets used, fuel spray tended to hit against the walls of the chamber and blow off in an im- perfectly burnt state. No exhaustive comparison was made of the two Merent arrangements but later tests have shown that there is not a great deal of difference between upstream and downstream injection and either may be used according to con- venience. Controlled atomization is, however, definitely preferable to relying for atomization on the penetration of a solid upstream jet.

For ignition we relied at first on a Diesel engine glow plug which was later modified and improved by having a priming jet built in, thus allowing fuel to be injected on to a wick sur- rounding the heater element. This worked quite well but took too much current and was later abandoned in favour of the spark plug now in general use.

In the B.10 unit there was no serious combustion problem since ample space was available, and it was not until the F.2 engine came into being that combustion really became a limiting factor.

In this engine, with its two-bearing shaft, it was necessary to limit the combustion chamber length not only to save weight but also to avoid running into the main whrling speed. I t was also desirable to keep the cross-sectional area of the chamber low in order to maintain a small frontal area.

When the engine was being designed, I had either to base the chamber design on the knowledge available at that time and accept a rather clumsy and bulky layout or else to gamble on future developments. I chose the latter course and allocated for combustion a space smaller than that considered necessary at that time, hoping that during the design and construction of the engine combustion development would make sufficient progress for combustion to be achieved in the space available. I think that with a normal type of combustion system my hopes would have been realized, but with the annular form actually used development is necessarily slow. Consequently, by the time the combustion design had to be completed the combustion development was still unsatisfactory, and it has had a retarding effect on the progress of the engine ever since.

The annular type of combustion system used in this engine is of considerable interest, and a typical design is shown in Fig. 19. By allowing air to flow through,a complete annulus instead of through a number of separate pipes, a larger area of flow for a given overall diameter can be achieved. The price that is paid, however, is that the chamber has to lje developed as a whole, whereas with the pipe system development can be carried out on a single pipe. The result is that the rate of development of an annular chamber is much slower than that of a system having separate pipes.

The result of the failure to estimate the probable rate of combustion development, which prejudiced the early days of the F.2 engine, was that much time was lost in trying to overcome the difficulties resulting from a combustion intensity higher than could be controlled by the techniques then available.

Right from the start of this work we had abandoned the steam turbine practice of dividing mbines into

Turbine Design.

two types-impulse and rea&on. This was, ofcourse, an in- evitable result of our conception of free vortex blading in which the degree of reaction increases progressively from root to tip.

1 We were faced, therefore, not with the alternative of designing


EARLY HISTORY OF THE A X I A L TYPE

Fig. 24b. F.2 Jet Propulsion Engine, Another View

Fig. 24c. F.2 Jet Propulsion Engine, Another View

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Plate 4 DEVELOPMENT OF THE I N T E R N A L COMBUSTION T U R B I N E

Fig. 24d. F.2 Jet Propulsion Engine, Another View

Fig. 260. Gloster F.9140 Aircraft with .F.2 Engines [I.Mech.E., 19451


423

Fig. 19. Arrangement of F.2 Combustion Chamber

for impulse or for reaction, but simply of deciding how much reaction to put in at the design radius.

There were at first very few reliable data to go on and our early decisions were based on little more than guess work. We were sure that with blades of 50 per cent reaction at the mean height (varying from about 40 per cent reaction at root to 60 per cent at tip) we ought to be able to attain a small-stage efficiency of about 90 per cent based on total head pressures.

We believed that if the degree of reaction was reduced so that the blades were pure impulse at the root and 40 to 50 per cent reaction at the tip, we should only be able to attain about 85 per cent efficiency. This loss of 5 per cent in turbine efficiency would make a considerable difference to the engines overall efficiency and power.

On the other hand there were a number of distinct advantages to be gained by reducing the amount of reaction. In the first place the temperature drop in the nozzles would be greater, thus leading to lower turbine blade temperatures. Again the heat drop per stage would be increased, leading to a lighter and more compact turbine. In spite of these points we decided that the balance of advantage lay with the high reaction blading and this was used in all the early designs.

I t soon became apparent, however, that, provided there was no actual recompression at the blade root so that the degree of reaction became negative, the loss in efficiency was very much less than the 5 per cent we had assumed. It appeared that the blading losses were not greatly affected by the actual amount of reaction, provided that the flow had some acceleration at all radii. In the next engine, the F. type, we accordingly designed for a considerable decrease in the amount of reaction. The natural consequence of t h i s was a change from a multistage drum type rotor to a single- or two-stage disk wheel.

Materials. Most of our compressors were made in light alloy, the material for the blades being RR.56. This material proved quite satisfactory and gave little trouble. After various methods of producing compressor blades had been tried a press- ing process was perfected by High Duty Alloys, Ltd., which produced a very cheap and accurate blade.

In Rex 78, Firth-Vickers produced an austenitic heat- resisting alloy which was adequate for all our requirements at the time. In the forged and heat-treated form it would with- stand a stress of nearly 3-0 tons per sq. in. at 750 deg. C. (1,382 deg. F.) for a creep strain of 0.1 per cent in 300 hours (which was our design requirement), and it had an ultimate tensile srrength of 22 tons per sq. in.

But the difficulty with the heat-resisting materials was not so much one of obtaining an alloy which had sufficiently good physical properties in the test piece as of getting sound material in the shapes needed for disks and drums. The steel manu- facturers had many problems before they overcame these di&ulties. We naturally dissipated a lot of effort in toying with unconventional materials, plastics for compressor blades, ceramics for turbine blades, etc., but could find no reasons for going ahead with any of them.

The F.2 Turbine Engine. We must now return to 1939 and trace the history of the jet propulsion unit, for which the compressor Freda was designed.

Immediately the war broke out, it became necessary for us to reconsider the whole of our gas turbine programme. Until that moment, we had been proceeding with a research objective of demonstrating the practicability of an aircraft gas turbine for propeller drive. We were aware that the complication of the power turbine made the project of a longer term character than that of the simple jet propulsion engine on which Power Jets were working. But we believed that for long-range aircraft, flying at the speeds that were in view at that time, there would be a need for a propeller turbine.

However, the urgency of war made it desirable to concentrate our efforts on projects which could be completed quickly. In September 1939 we therefore suggested to Power Jets that a jet propulsion engine should be constructed on the basis of the D.l l design, the power turbine being omitted.

My original conception of the F. type jet propulsion engine is shown in Fig. 20. This design, known as the F.1, was produced in December 1939 and it provided for a unit giving 2,150 lb. static thrust, a pressure ratio of 4/1, a maximum temperature of 800 deg. C. (1,472 deg. F.) with a mass flow of 38.0 lb. per sec. The design speed was 9,450 r.p.m., the overall diameter 27 inches, and the length 7 ft. 9+ in. The compressor was of nine stages with a rotor having a disk wheel construction; the combustion chamber was of annular, straight-through layout; the compressor was driven with a single-stage water-cooled turbine; provision was to be made for control of the compressor boundary layers by air bleeds; and bearing lubrication was by oil bath with no circulating system.

We had hoped to get th is engine manufactured as a joint effort by Power Jets and R.A.E., and plans were made to this effect. In July 1940, however, Power Jets had to withdraw owhg to the pressure of other commitments, and the work was taken over by the Metropolitan-Vickers Company instead, but by the time t h i s change was made the design had evolved considerably. In the new design, the F.lA, the capacity had been increased to 47.5 lb. per sec. giving 2,690 lb. of static thrust at 7,470 r.p.m. As wi l l be seen from Fig. 21, the compressor disk wheels had been replaced by a drum, the water-cooled single-stage turbine had become air cooled with two stages, and the intermediate bearing had disappeared.

The change to the drum was, I now believe, a retrograde step, occasioned by our fears of losing the blade tip clearance owing to unequal cooling of rotor and stator. We changed to a two-stage turbine in order to reduce the overall diameter of the unit. Although we achieved our object, the net result was probably in the wrong direction, for it involved an increase in weight and may have added starting difliculties. The loss of the centre bearing increased the weight and later introduced whirling speed difficulties. I must admit that to me the two-bearing arrangement still looks right and I have been unable to reconcile th is with the better performance on paper of the three-bearing system.


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EARLY HISTORY O F THE AXIAL TYPE

Fig. 266. Gloster F.9/40 Aircraft with F.2 Engines, Another View

Fig. 26c. Gloster F.9/40 Aircraft with F.2 Engines, Another View

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426 DEVELOPMENT O F THE INTERNAL COMBUSTION TURBINE There is, however, no doubt that the introduction of air

cooling was a beneficial change. Its adoption was due to the success resulting from a similar change in the contemporary Power Jets engine. On the whole, the first six months of 1940 was a period during which inspiration was conspicuous by its absence. Perhaps this may have been due to the unsettling effects of international friction.

It was this design, together with all existing material, which was handed over to Metropolitan-Vickers in July 1940. The layout of the engine (F.2) as finally manufactured, tested and flown, is shown in Fig. 22 and photographs of the compressor rotor and complete engine are shown in Figs. 23 and 24, Plates 2, 3 and 4.

The F.2 engine ran for the first time in December 1941, and test results are shown in Fig. 25. The leading particulars of the engine are given below :-

Static thrust . . 2,200 lb. Specific fuel consumption . 1.07 lb. per hour per lb. Maximum speed . . 7,390 r.p.m. Maximum temperature . 800 deg. C . (1,472 deg. F.) Weight . . 1,530 lb. Overall diameter . . 36inches Overall length . . 103 feet.

It may be mentioned here that the engine was superior to its German contemporary the Jumo 004 by 25 per cent in respect of specific fuel consumption and by 5 per cent in respect of specific weight. The engine flew for the first time in the F.9/40 fighter aircraft in November 1943. The aircraft with these engines installed is shown in Fig. 26, Plates 4 and 5.

The generalized performance curves for jet propulsion engines at 500 m.p.h. in the stratosphere are shown in Fig. 27. The design performance of the F.2 engine is approximately indicated on this Fig. The big advances that have yet to be made by increase in pressure ratio and maximum temperature should be noted.

By this time the F.2 engine had shown what it could do in its original form, and further work on it became largely a matter of development. The responsibility for this lay with the engineers of the Metropolitan-Vickers Company and I will leave this part of the story to them. This is a convenient opportunity for me to pay a tribute to the tirelessly thorough work of the engineers of this company, with whom I had the pleasure and privilege of working during these early days. My own interests and duty lay with research, and having passed this engine on to Metropolitan-Vickers, I had to leave it in their capable hands and turn my attention to the next step forward.

Perhaps forward is the wrong word to use here. For it will not have escaped your notice that most of our time during the early years was devoted to attempts to retrieve errors made by depart- ing from the conceptions of 1936. If, in the fullness of time, the wheel is to go full circle, we may return once more to the original idea of the double compound axial engine with coaxial shafts, similar to that shown in Fig. 7-similar, but I hope with a wealth of detail differences.

There has not been space here to do more than outline the more important events that occurred in the early history of the axial type of gas turbine engine in this country. The record has been carried from 1936 (when the work effectively started) to about 1942 (when the F.2 engine passed from the research to the development stage).

In parallel with the engine work, a considerable amount of research was proceeding on the related problems of combustion, aerodynamics, thermodynamics, and stressing. But al l that is another story.

In the five years covered by this review, we succeeded in producing an axial compressor with an overall efficiency of 84 per cent at a pressure ratio of 6/1, a multistage turbine with an overall efficiency of 89 per cent, a jet propulsion turbine engine with a thermal efficiency of 22 per cent, and a lot of ideas for the future. Some of these ideas have already borne fruit and I hope that more will do so in the days to come.

Conclusions.


Documents

The Early History of the Axial Type of Gas Turbine Engine