Method of and Apparatus For Improved Utilization of the Thermal Energy Contained in a Gaseous Medium

Abstract

The present invention concerns a method of utilising the waste heat contained in the exhaust gas of an internal combustion engine, comprising a turbine (20). To provide an apparatus and a method of operating same which directly supplies additional drive energy which otherwise would be lost as waste heat, it is proposed according to the invention that the turbine is an inverse turbine connected downstream of the exhaust gas outlet of the internal combustion engine and comprising at the inlet side an expansion stage (23) and at the outlet side a subsequent compressor (21), wherein the expansion stage and the compressor of the inverse turbine are so designed that the downstream-disposed compressor of the inverse turbine generates at the outlet of the expansion stage (23) a reduced pressure (p1) below the ambient pressure (p0), wherein the outlet (2b) of the compressor (21) is at the level of the ambient pressure and the compressor of the inverse turbine is driven by the turbine.

Claims

1. A method of utilising the waste heat contained in the exhaust gas of an internal combustion engine, comprising a turbine (10), characterised in that at least a part of the exhaust gas acts on an inverse turbine (20) which at the inlet side comprises an expansion stage (23) and at the outlet side a subsequent compressor (21), wherein the expansion stage and the compressor of the inverse turbine are operated in such a way that the downstream-connected compressor of the inverse turbine generates at the outlet of the expansion stage (23) a reduced pressure (p1) below the ambient pressure (p0), wherein the outlet (2b) of the compressor (21) is at the level of the ambient pressure and the compressor is driven by the turbine.

2. A method according to claim 1 characterised in that it is applied to a gas turbine, wherein the pressure ratio of the compressor of the gas turbine is set to at least 10, and the pressure ratio between the outlet and the inlet of the expansion stage is set to at least 10.

3. A method according to claim 1 characterised in that the reduced pressure at the outlet of the expansion stage of the inverse turbine is promoted by intercooling, i. e. by means of an intercooler, in particular a heat exchanger.

4. A method according to claim 1 characterised in that the inverse turbine is operated at a speed of rotation which is different from the speed of rotation of the upstream-disposed turbine.

5. Apparatus for utilising the waste heat contained in the exhaust gas of an internal combustion engine, comprising a turbine (11, 13), characterised in that the turbine is an inverse turbine connected downstream of the exhaust gas outlet of the internal combustion engine and comprising at the inlet side an expansion stage (23) and at the outlet side a subsequent compressor (21), wherein the expansion stage and the compressor of the inverse turbine are so designed that the downstream-disposed compressor of the inverse turbine generates at the outlet of the expansion stage (23) a reduced pressure (p1) below the ambient pressure (p0), wherein the outlet (2b) of the compressor (21) is at the level of the ambient pressure and the compressor of the inverse turbine is driven by the turbine.

6. Apparatus according to claim 5 characterised in that the expansion stage (1) and the compressor (2) rotate about the same axis.

7. Apparatus according to claim 5 characterised in that the compressor (2) is connected to the turbine shaft by way of an interposed transmission (5).

8. Apparatus according to claim 5 characterised in that the volume of the compressor (11) and the expansion chamber (13) of the turbine is respectively smaller by a factor of 1.2 to 4, than the volume of the compressor (21) and the expansion stage (23) of the inverse turbine.

9. Apparatus according to claim 5 characterised in that the expansion stage and/or the compressor have at least one diffuser, the interior of which has a coolant at least partially flowing therethrough.

10. Apparatus according to 9 claim 5 characterised in that the expansion stage (1) is connected downstream of the last stage of a gas turbine (6) and is driven by the exhaust gas from the gas turbine (6) and the pressure drop generated by the compressor (2).

11. A combination of a gas turbine (19) with an inverse turbine (20) according to claim 5 characterised in that devices for the feed of additional fuel are provided after the outlet of the gas turbine and preferably before or in the inverse turbine.

12. Apparatus according to claim 5 characterised in that the expansion stage (1) is connected to an exhaust gas outlet of an internal combustion engine and is driven by the exhaust gas of the internal combustion engine and the pressure drop generated by the compressor (2).

13. An aircraft engine characterised by an apparatus according to claim 5.

14. An aircraft engine according to claim 13 comprising an outer housing 1 and an inner engine housing 2, between which there are provided support elements and a stationary diffuser (26), wherein a fan (25) driven by a first turbine shaft (14) generates an axial air flow between the outer (1) and the inner housings (2), characterised in that provided at the end of the turbine shaft (14), that is remote from the fan, is an inverse turbine (23b) with a concluding compressor (21) and a cooling device (22) arranged therebetween.

15. An aircraft engine according to claim 13 comprising a heat exchanger between the expansion stage and the compressor of an inverse turbine, wherein cooling conduits extend through the housing and possibly through diffusers provided in the cooling device and in diffuser blades and contain a coolant to cool exhaust gases in the cooling chamber (22).

16. A method according to claim 1 characterised in that it is applied to a gas turbine, wherein the pressure ratio of the compressor of the gas turbine is set to at least 15, and the pressure ratio between the outlet and the inlet of the expansion stage is set to at least 15.

17. Apparatus according to claim 6 characterised in that the expansion stage (1) and the compressor (2) are mounted on a common shaft.

18. Apparatus according to claim 8 characterised in that the volume of the compressor (11) and the expansion chamber (13) of the turbine is respectively smaller by a factor of 1.2 to 2.5 than the volume of the compressor (21) and the expansion stage (23) of the inverse turbine.

Description

[0033] Further advantages, features and possible uses of the present invention will be clearly apparent from the description hereinafter of a preferred embodiment and the related Figures.

[0034] FIG. 1 diagrammatically shows the structure of a conventional gas turbine with compressor, combustion chamber and turbine rotor,

[0035] FIG. 2 shows a diagrammatic view of the conventional gas turbine of FIG. 1 together with a downstream-connected inverse turbine,

[0036] FIG. 3 shows a thermodynamic TS diagram relating to the combination of FIG. 2,

[0037] FIG. 4 shows a diagrammatic view of a combination similar to the combination of FIG. 2 but with a further combustion chamber disposed upstream of the inverse turbine for complete combustion of exhaust gases, and

[0038] FIG. 5 shows the structure of an aircraft engine with downstream-connected inverse turbine.

[0039] FIG. 1 firstly shows the diagrammatic layout of a conventional gas turbine 1 which substantially comprises a compressor 11, a combustion chamber (not shown) and the turbine stage 13 and which are mounted in a suitable turbine housing for example on a common shaft.

[0040] A gaseous fuel or spray mist of fuel is injected into and fired in the combustion chamber so that the combustion gases drive the turbine stage 13. That in turn is carried on a common shaft 4 with the upstream-connected compressor 1 and drives same which thereby compresses additional combustion air and forces it into the combustion chamber.

[0041] The corresponding inverse turbine 20 which according to the invention could be connected downstream of the turbine 10 in FIG. 1 is shown in FIG. 2 and at the inlet side comprises an expansion stage 23, followed by an intercooler 22, and then a compressor 21.

[0042] The only diagrammatically illustrated intercooler can be implemented for example in the form of a heat exchanger, for example by cooling conduits which extend in the interior of one or more diffusers and an expansion stage or compressor housing.

[0043] The hot gas issuing from the last stage 13 of the turbine 10 is fed to the expansion stage 23 of the inverse turbine 20 which is driven thereby and which in turn by way of a common shaft 24 drives the compressor 21, the inlet side of which is connected to the outlet side of the expansion stage 23. As a result a reduced pressure is generated at the outlet of the expansion stage 23 so that the increased pressure drop causes an increase in power of the rotor of the expansion stage 23. A part of that additional power is used by the downstream-connected compressor 21 which however at a lower temperature level and with a lower pressure drop or a lesser pressure rise compresses the exhaust gas again to the ambient pressure and accordingly has a compressor outlet 21b open to the environment. The lower temperature level achieved by the intercooler 22 allows an increase in the pressure drop.

[0044] FIG. 3 shows the additional energy gain idealised in a schematic temperature-entropy diagram (TS diagram).

[0045] Adiabatic compression by a compressor between points 1 and 2 in the TS diagram is followed by the feed of heat and the increase in pressure and temperature from the point 2 to the point 3. The following adiabatic relief would end at the point 4 without the downstream-connected inverse turbine, whereupon once again cooling would be effected with a reduction in pressure and temperature towards point 1.

[0046] The inverse turbine or the downstream-connected compressor of the inverse turbine allows greater adiabatic expansion at the point 5, from where in turn gas is cooled along the path from point 5 to point 6 and then is raised by the compressor from the point 6 to a somewhat higher temperature at point 7, from where the cycle can begin again at 1.

[0047] In that way an additional amount of energy which in the TS diagram corresponds to the area of the rectangle defined by the points 4, 5, 6 and 7 is obtained.

[0048] FIG. 2 shows the diagrammatic layout of an inverse turbine combined with a conventional turbine, wherein the inverse turbine, in the case of the typically lesser pressure and temperature drop in the exhaust gas, is of greater dimensions. That effect however is in part also compensated by the lower temperature of the exhaust gases so that in practice the inverse turbine often does not have to be of larger dimensions, or of only slightly larger dimensions, than the upstream-connected conventional gas turbine. Alternatively the speed of rotation of the inverse turbine 11 can be increased in relation to the speed of rotation of the upstream-connected gas turbine 10, for example by an interposed transmission, and, with a suitable design configuration, the speed of rotation of the compressor 21 can differ from that of the expansion stage 23.

[0049] A further example is a combination with a combustion chamber additionally provided between the conventional gas turbine and the inverse turbine, as shown in FIG. 4, wherein that variant is referred to herein as a “crossover turbine”, and wherein the additional combustion chamber is intended in particular to serve for more complete combustion of the exhaust gases from the combustion chamber of the first turbine, but in addition can also be supplied with further fuels. Because of the higher temperatures that permits a marked increase in power even if the dimensions of the inverse turbine are possibly to be correspondingly increased. Desirably the inverse turbine is mounted on a shaft 24 which is separate from the shaft 14 of the gas turbine in order generally to differently set the speeds of rotation of the conventional turbine and the inverse turbine, in consideration of the prevailing pressure, temperature and flow conditions.

[0050] FIG. 5 is a view which is admittedly still diagrammatic but rather closer to reality than the other Figures of essential components of an aircraft engine 100 with a downstream-connected inverse turbine.

[0051] The aircraft engine 100 shown in FIG. 5 has an outer housing 1 and an inner housing 2 which are connected together by way of a diffuser 21 which comprises guide blades 26a which are distributed over the periphery and which join the inner housing 2 and the outer housing 1 together in the manner of spokes. A first shaft 14 carrying a plurality of different blade rings 111, 123 and 122 is rotatably mounted in the inner housing 2. In addition, mounted on the first shaft 14 in a recessed portion is a second shaft 24 which also carries a part of the rings 111 and blade rings 113. Two diagrammatically illustrated bearings 19 illustrate the rotatability of the shaft 24 with respect to the shaft 14. The rotatable mounting of the shaft 14 to the housing 2 and/or 1 is not shown here. One or more blade rings which are arranged in part on the shaft 14 and in part on the shaft 24 respectively belong to different compressor and turbine stages.

[0052] Arranged at the front end of the shaft 14 is a so-called fan 25, that is to say a blade ring of large diameter, which substantially fills up the inside diameter of the outer housing 1. The fan is driven by the first shaft 14. The diffuser comprises a ring of guide blades which however are arranged rigidly between the housing 1 and the inner housing 2 and which serve to orient the air flow generated by the fan optimally along the axis to produce a maximum amount of thrust.

[0053] The mounting of the shaft 14 in the inner housing is not shown here but can also be implemented for example in the region of a cooling chamber 22 and at the front end of the shaft 14 behind the fan and at the inner housing 2 respectively.

[0054] At least a part of the blades 26a have inner cooling passages 27 which serve to cool a coolant in the interior of the diffuser blades which is passed by way of the inner housing 2 into the region of the cooling chamber 22 and in addition can possibly also circulate through rigid guide blades or spokes in the region of the cooling chamber 22. The flow of coolant is otherwise only diagrammatically shown in FIG. 5 outside the housing 2.

[0055] The individual components which occur in succession along the axis can respectively each have more or fewer blade rings than shown here and the axial length of the engine in relation to the diameter is here not necessarily reproduced in the correct relationship.

[0056] Substantially the engine comprises a low pressure compressor 11a formed by blade rings 111 on the first shaft 14, a subsequent high pressure compressor 11b which is formed by similar blade rings 111 on the second shaft 24, a combustion chamber 12a which follows the high pressure compressor 11b, a high pressure turbine portion 13a equipped with blade rings 113 and finally followed by the low pressure turbine 23a. The low pressure turbine 23a would form the axially rearward end of a conventional aircraft engine, wherein, as stated, the number of blade rings 123 could also be greater and the low pressure turbine portion could be longer, so that for example it would also embrace the region 23b.

[0057] The regions 23a and 23b however are here deliberately shown separately, wherein the portion 23a is attributed to the low pressure turbine of the conventional engine while the region 23b is to be viewed as the turbine portion of an inverse turbine which is finally also followed by a compressor 20 with blade rings 121. Disposed between the outlet of the low pressure turbine 23a and the inlet of the inverse turbine 23b is a further combustion chamber 15 which serves for complete combustion of constituents which hitherto have not been burnt. Optionally additional fuel could also be injected here. Disposed in the turbine portion 23b of the inverse turbine and the compressor 21 of the inverse turbine there is also a cooling chamber serving to further cool the exhaust gases from the inverse turbine 23b, which have already been considerably relieved of pressure but which are still hot, which in addition to the action of the compressor 21 generates a reduced pressure after the turbine portion 23b and makes it possible to maintain the gas flow with a lower level of compression power.

[0058] The blade rings 111, 123 and 121 on the first shaft 14 are designed for low peripheral speeds while the blade rings 111 and 113 on the second shaft 24 which can be mounted on the inner shaft 14 or however in the region of the combustion chamber 12a involve a higher peripheral speed.

[0059] The flow generated by the high pressure turbine however by way of the blade rings 123 drives the first shaft 14 which in turn drives the fan which generates the main part of the overall thrust.

[0060] With the additional relief of pressure of the working gas in the expansion stage 23b of the inverse turbine and subsequent compression additional thrust is generated, which increases the efficiency of the engine by some percent and correspondingly reduces the fuel consumption.

LIST OF REFERENCES

[0061] 1 outer housing [0062] 2 combustion chamber [0063] 3 turbine rotor [0064] 4 shaft [0065] 6 gas turbine [0066] 10 gas turbine [0067] 11 turbine rotor [0068] 11a low pressure compressor [0069] 11b high pressure compressor [0070] 12a combustion chamber [0071] 13 compressor [0072] 13a high pressure turbine portion [0073] 14 shaft [0074] 15 combustion chamber [0075] 16 first shaft [0076] 19 gas turbine [0077] 20 inverse turbine [0078] 21 compressor [0079] 21b compressor outlet [0080] 22 intercooler [0081] 23 expansion stage [0082] 23a low pressure turbine [0083] 23b turbine [0084] 24 second shaft [0085] 25 fan [0086] 26 diffuser [0087] 26a guide blades [0088] 27 passages [0089] 100 aircraft engine [0090] 111 running ring; blade ring [0091] 113 running ring; blade ring [0092] 121 running ring; blade ring [0093] 122 blade ring [0094] 123 running ring; blade ring