US2435138A - Compound internal-combustion turbine plant - Google Patents

Compound internal-combustion turbine plant Download PDF

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US2435138A
US2435138A US517272A US51727244A US2435138A US 2435138 A US2435138 A US 2435138A US 517272 A US517272 A US 517272A US 51727244 A US51727244 A US 51727244A US 2435138 A US2435138 A US 2435138A
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Prior art keywords
blades
turbine
compressor
shell
pressure
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Expired - Lifetime
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US517272A
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Heppner Fritz Albert Max
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Armstrong Siddeley Motors Ltd
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Armstrong Siddeley Motors Ltd
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/06Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
    • F02C3/073Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages the compressor and turbine stages being concentric

Description

Jan. 27, 1948. F. A. M. HEPPNER 2,435,138

COMPOUND INTERNAL- COMBUS T ION TURB INE PLANT Filed Jan. 6, 1944 7 Sheets-Sheet 1 7 )4 [Ma-wrap JTTOP/VEX' Jan. 27, 1948. F. M. HEPPNER 2,435,138

COMPOUND INTERNAL-COMBUSTION TURBINE PLANT Filed Jan. 6, 1944 7 Sheets-Sheet 2 INVENTOR Z. z r M Arrmrni's'i Jan. 27, 1948. F. A. M. HEPPNER 2,435,138

COMPOUND INTERNAL-COMBUSTION TURBINE PLANT I Filed Jan. 6, 1944 '7 Sheets-Sheet 3 14.5 W MW flrro ENE) Jan. 27, 1948. F. A. M. HEPPNER COMPOUND INTERNAL-COMBUSTION TURBINE PLANT 7 Sheets-Sheet 4 Filed Jan. 6, 1944 .a oz- AM! Jim ATTORNEY Jan. 27, 1948. F. A. M. HEPPNER 2,435,133

COMPOUND INTERNAL-COMBUSTION TURBINE PLANT Filed Jan. 6, 1944 7 sheets-sheet e Disc-m goNDmoNs.

VOLUME.

REGULATDN CONDITIONS.

m z E/vrofi' VOLUME. W m

W Mm flTTO/YIVE) Jan. 2?, 1948. F. A. M. HEPPNER 2,435,138

COMPOUND INTERNAL-COMBUSTION TURBINE PLANT Filed Jan. 6, 1944 '7 Sheets-Sheet 7 DiRECTION OF AIR STATIONARY l NOZZLE DIRECT now 20 OF ROTATIGN IMPULSE REACTION REACTION w as mnomu.

3n van for Gtl'onuag Patented Jan. 27, 1948 COMPOUND INTERNAL-COMBUSTION TURBINE PLANT Fritz Albert Max Hemmer, Leamlngton SpmEng- Armstrong land, assignor to Siddeley Motors Limited, Coventry, Englan Application January 8, 1944, Serial No. 517,272

In Great Britain January 15, 194 3 Claims. ,(Cl. 230-122) This invention relates to a compound, intemals purposes, the contra-rotating shell carrying externally jet-augmenter blades.

The specification of my co-pending patent application Serial No. 500,694, filed August 31, 1943, now matured into Patent No. 2,426,098, of August 19, 1947, discloses such a compound plant in which useful work is taken out between adjacent turbine sections, so that the pressure at the end of a compressor section is substantialfy equal to that at the beginning of the associated turbine section, whereby I avoid sealing problems, particularly at large diameters.

In the specification of my co-pending patent application Serial No. 500,695, filed August 31, 1943, now abandoned, the contra-rotating turbine shell coacts with independent turbine sections respectively coupled to compressor sections, and useful work is taken out of the gases and transmitted to the shell between most orall of the turbine sections, the shell being driven relatively slowly.

For perfect pressure balance between the turbine and compressor sections the speed should be uniform, i. e., all the sections should rotate at the same rotatable speed if the blade efficiency is constant throughout, but this raises difiiculties in the design of the compressor, and regulation is entirely unsatisfactory.

It is my main object to avoid these disadvantages.

A further object is to design the shell to run more slowly than it would on the reaction of the compressor only, as in my applications above mentioned.

For an understanding of these and other objects and advantages of the invention attention should be directed to the following descrippanying diagrammatic drawings in which- 2 of and Figure 5 a fragmentary developed plan of the independent stage;

Figure 6 shows a temperature-entropy diagram 01' the turbine in design conditions, and Figure 6a in regulation conditions;

Figures '1 and 7a show corresponding pressure-volume diagrams; and

'Fig. 8 is a diagrammatic illustration of the structure and arrangement of the turbine blading oi. Fig. 2.

According to one feature of my invention, a compound internal-combustion turbine plant comprises a number of independent compound rotors each carrying compressor and turbine blades, the latter coacting with blades on a contra-rotating shell, the compressor blades coacting with stationary compressor blades, and useful work is taken out of the contra-rotating shell between the various rotors, and additional useful work is taken out at the high-pressure end by reaction on stationary blades, the various rotors having higher rotational speeds as the volume decreases to achieve practically constant incidence of the whole blading under regulation and starting conditions.

It may here be stated that if the shell were not driven against the said other stationary blades the speed with which the shell would rotate would be determined by the compressor torque and the total amount of available useful work, because, it a certain compressor torque a is available and a certain amount 2) of useful work is available,

. .the useful work can only be extracted if aXnXK Figure 1 is a graph of the operation of a com- I equals b, where n is the speed and K is a con stant. Therefore,

If the shell is to be driven more slowly (i. e.,

according to the invention) additional torque must be provided, whence where a is the additional turbine torque.

Thus, the additional torque necessary to drive the shell may be provided by arranging for the gases from the combustion chamber to pass at high speed through stationary nozzles to the first row or rows (which is or are on the shell) of turbine blades, these blades being partly or wholly of the impulse type, whilst the next two rows of turbine blades (at least) are of the reaction type, the final rows (which are not on the shell) being independently rotatable with their associated 3 compressor rows and being aerofoil blades-i. e., well-spaced blades.

The result is that the turbine operates with optimum efliciency, inasmuch as losses at the high-pressure end are to a great extent regained in the low-pressure end of the turbine, which latter end is very eflicient.

Referring now to the drawings, the graph of Figure 1 is of a turbine plant according to the specification first aforesaid, and that of Figure 1a of the plant shown in Figure 2. The vertical lines A represent independent turbine and compressor sections, the ordinates P representing pressure. The areas enclosed by the lines B, C represent unavoidable pressure-difierence, the line B being for the compressor and the line C for the turbine in both graphs. With the arrangement of the invention the resultant area enclosed between the lines C and B is very much reduced, it will be observed.

Referring now to Figures 2 to 5, the plant illustrated is one in which the turbine H is disposed radially outwardly of the compressor 22, the flow through the latter being from the top to the bottom. The compressor comprises stationary blade rows l3, l3 and rotary rows 54, l i, the latter rows being fast with independent turbine rows i5, IS. The first independent turbine row 15a is mounted to revolve with a sleeve is carrying a plurality of compressor rows Ida coacting with stationary compressor rows Mia.

The burning gases are led from the tubular members it to a stationary row of blades 18 forming nozzles through which the gases pass at high speed to the first row 2d of turbine blades mounted on the shell 2 I, these blades being partly or wholly of the impulse type. (The shell carries external augmenter blades 22 coacting with stationary augmenter blades 23.) The row of turbine blades 15a fast on the sleeve l6 rotates at high speed to drive the final highest pressure portion of the compressor, this row and the next row 24 (on the shelz) being of the reaction type. The final independent rows and coacting rows 25 on the shell are all of the aerofoil type. The arrangement and characteristics of the blades i9, 20, I5a, 24 and I5 are indicated diagrammatically in Fig. 8.

As will be seen from Figures 3, 4 and 5, the blades I 5 are spaced from one another sufliciently to allow of assembly into the shell through the larger (upper) end thereof, the blades passing between the blades 25 fast with the shell. Figure 5 shows at Mb some of the compressor blades strengthened to carry the centrifugal forces.

Conveniently the shell 2| is mainly of aluminium, the blades 25 being riveted to sheetmetal, arcuate portions 21 having end clearance from one another when cold, as indicated at 28 (Figure 4), the portions 21 being riveted at 29 to aluminium ribs 30 with heat insulating material 3| interposed. Alternatively, the sheil can be made from steel in one piece with the turbine blades 25 permanently fixed in it, this being possible by the new way of assembly as a consequence of the wide spacing of the turbine blades. This feature also allows of assembling and dismantling the shell without interfering with the balance, which is a very important point because of the relatively-great weight of the shell.

At 33 I show a drive shaft, for auxiliaries, passing through the hollow feed nozzle 34, the drive shaft carrying a gear 35 which is driven from the compressor sleeve l6 through a compound pinion 35 and gear 31.

As previously stated. perfect pressure balance asks for equal revolutionary speeds of all the compound members, but this interferes with re ulation, i. e., operation at a reduced fuel supp y. and it also interferes with constant Mach numher on the compressor, because the rotational speed remains constant but the Mach number goes up with temperature. Furthermore, as the volume decreases higher revolutionary speeds and smaller dimensions are required to keep the wheels at their optimum proportions. The optimum compressor shows constant Mach number to reduce the number of stages. The replacement of a great number of counter-rotating discs by one shell gives the mechanical mechanism an aerodynamic rigidity, which must be made flexible to make the incidence constant on the blades in regulation conditions. This is done, according to the invention, by thermodynamic means, taking out a certain amount of useful work at the high pressure end of the counter-rotating shell against stationary turbine blades so that the high pressure compressor part is driven at a turbine pressure which is a certain amount lower than the corresponding compressor pressure. Then, in regulation conditions, the available work for driving the high pressure compressor becomes smal.er because the diagram (see Figures 6a and 7a) must read as though the lines of constant pressure become more closely spaced towards the low pressure region. Therefore overspeeding of the high pressure part at partial loads is avoided, and the incidence on all the blades (turbine as well as compressor) can be kept practically constant throughout the whole range of regulation.

Additional practical advantages are- (1) Constant Mach number in compressor.

(2) High temperature of turbine on slowlyrunning small-diameter sheil part running under low stresses.

Referring to Figure 6, or Figure 6a, AB represents the rise in temperature in the compressor, assuming adiabatic compression, and CD the temperature drop in the turbine on the assumption of adiabatic expansion, whilst BC is a constant pressure line representing the heat taken in during combustion. ATI and AT6, on the line AB, represent equal temperature rises (i. e., equal increments of heat) in the first six compressor sections, and AT] the temperature rise in the last (highest-pressure) section-assuming seven sections altogether, while ATI to AT? on the line CD are the corresponding temperature drops for the turbine. The useful work removed at X represents the energy bleed against the stationary blades l9, i. e., the work done on the first shell blades 20. The dotted lines are constant pressure lines which are, however, given a droop adjacent the line CD pictorially to represent the modification introduced by the invention. In a similar way the dotted lines in Figures 7 and 7a are intended pictorially to indicate the alteration, due to the invention, in the P/V diagrams.

With regard to Figures 6 and 6a, consider the line Bit-C6. The correspondin iine B4C4 in Figure 2 of my application aforesaid No. 500,694 is purely a constant-pressure line. In the present instance, however, the shell speed is lower. Therefore the amount of useful energy actually taken out between the first two compound discs Ma: and, My (see Figure 2) is less than that corresponding to the thermodynamic efficiency of that section of the engine. This accounts for the "droop" in the line 36-06. For the same reason there is a "droop in the line 35-05 which is greater than that in Bi-CO, the corresponding compound discs being those marked My and 2. And so on. In other words, the turbine portion of a compound section works at a lower pressure than the corresponding com-,

pressor portion. In consequence, in regulating down, i. e., at reduced fuel, the available work of the turbine portion decreases more quickly than would be the case if these droops did not exist, thus avoiding over-speeding 01' the highestpressure portion of the compressor. I further efiect a transfer of energy to turbine portions connected to low-pressure parts of the compressor which otherwise would be driven too slowly and would have negative incidence, giving low eiiiciency. When regulating down the speed fall in the lowest-pressure sections is greater than in the highest-pressure sections, as appears from a; comparison of Figures 6 and 6a. The loss or work represented by these droops is made up for by the energy bleed X.

Itwillbeunderstoodthatifonetooktheextra work X (Figures 6 and 60) out at the low-pressure end instead of at the high-pressure end, instead or "droops one would have rises," and regulation eflects would be opposite to those described, 1. e., would be impossible.

Incidentally, the combined "droop and rise" shown in the line A-D is accounted tor by the exhaust oi the engine reacting against the dirfuser blades.

Finally, it may be added that the term Mach num is here used in its ordinary significance, to denote, at a given temperature, the ratio of the speed of air (relatively to a blade) to the speed of sound. There may, of course, be difierent Mach numbers, for a given temperature, for the leading and trailing edges of a blade.

What I claim as my invention and desire to secure by Letters Patent of the United States is:

1. A compound internal-combustion turbine plant including a plurality of independent compound rotors each rotating in the same direction and carrying compressor and turbine blades, a shell rotating in the opposite direction to said rotors and carrying turbine blades coacting with said first-named turbine blades, said shell taking out useful work between said rotors, stationary compressor blades coacting with said first-named high-pressure end or the turbine blades. the adjacent turbine blades carried by the shell coacting with said stationary blades to take out additional useful work at the high pressure end of the turbine, said blades being so constructed and arranged that the rotors have higher rotational speeds as the volume of gases decreases.

2. A compound internal-combustion turbine plant according to claim 1 wherein the stationary blades are stationary nozzles cooperating with the first row of turbine blades on the shell at the high-pressure end of the turbine and through which the combustion gases pass at high speed to said blades to provide additional torque for driving the shell, said blades being at least in part or the impulse type, at least two of the next succeeding rows of turbine blades being of the reaction type. and the remainin rows of turbine blades on the rotors being aerofoil blades,

3. A compound internal-combustion turbine plant including a plurality of independent compound rotors rotating in the same direction and carrying compressor and turbine blades, stationary compressor blades coacting with said firstnamed compressor blades, an output member in the form or a shell rotating in the opposite irection to said rotors and carrying rows of turbine blades coacting with said first-named turbine blades, the rotors having higher rotational speeds as the volume of gases decreases, at least part of the highest pressure turbine blades on the shell being of the impulse type, and stationary nozzles enacting with said highest pressure turbine blades and directing the hot gases to said blades at high speed, the next succeeding row of turbine blades after said highest pressure blades being mounted on the highest pressure rotor and being of the reaction type, and the next succeeding row of turbine blades being also of the reaction type and being fast with said shell and with a row or compressor blades.

FRITZ ALBERT MAX HEPPNER.

file of this patent:

FOREIGN PATENTS Number Country Date 411,473 France Apr. 12, 1910 541,349 Great Britain Nov. 24, 1941

US517272A 1943-01-15 1944-01-06 Compound internal-combustion turbine plant Expired - Lifetime US2435138A (en)

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Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2497444A (en) * 1945-02-02 1950-02-14 Fairey Aviat Co Ltd Cooling means for aircraft engines
US2613501A (en) * 1945-06-02 1952-10-14 Lockheed Aircraft Corp Internal-combustion turbine power plant
US2785849A (en) * 1948-06-21 1957-03-19 Edward A Stalker Compressor employing radial diffusion
US2788172A (en) * 1951-12-06 1957-04-09 Stalker Dev Company Bladed structures for axial flow compressors
US2790596A (en) * 1953-08-06 1957-04-30 Leo M Stirling Dual fan construction

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR411473A (en) * 1910-01-11 1910-06-17 Emile Baptiste Merigoux Turbo-compressor
GB541349A (en) * 1940-07-19 1941-11-24 Bertram Tom Hewson Improvements in and relating to prime movers, more particularly gas turbines for the propulsion of aircraft

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR411473A (en) * 1910-01-11 1910-06-17 Emile Baptiste Merigoux Turbo-compressor
GB541349A (en) * 1940-07-19 1941-11-24 Bertram Tom Hewson Improvements in and relating to prime movers, more particularly gas turbines for the propulsion of aircraft

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2497444A (en) * 1945-02-02 1950-02-14 Fairey Aviat Co Ltd Cooling means for aircraft engines
US2613501A (en) * 1945-06-02 1952-10-14 Lockheed Aircraft Corp Internal-combustion turbine power plant
US2785849A (en) * 1948-06-21 1957-03-19 Edward A Stalker Compressor employing radial diffusion
US2788172A (en) * 1951-12-06 1957-04-09 Stalker Dev Company Bladed structures for axial flow compressors
US2790596A (en) * 1953-08-06 1957-04-30 Leo M Stirling Dual fan construction

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