GAS TURBINE

Abstract

A gas turbine includes a rotor that is rotatable about an axis, a casing configured to cover the rotor in a circumferential direction and having an annular space therein, a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing, a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas, a turbine driven by the combustion gas, and an air introduction passage defined by a partition plate configured to divide the space in the casing in the circumferential direction of the rotor and an inner circumferential surface of the casing and configured to introduce the compressed air in the casing into the combustor.

Description

CROSS-REFERENCE TO RELATED APPLICATION

Priority is claimed on Japanese Patent Application No. 2021-173894, filed Oct. 25, 2021, the content of which is incorporated herein by reference.

BACKGROUND OF THE INVENTION

Field of the Invention

The present disclosure relates to a gas turbine.

Description of Related Art

A gas turbine includes main components such as a compressor configured to compress air, a combustor configured to combust fuel using the air compressed by the compressor and generate a combustion gas, and a turbine driven by the combustion gas from the combustor.

The combustor has an outer shell and a combustor basket provided in the outer shell, and an air flow channel into which the compressed air compressed by the compressor flows is formed between an inner circumferential surface of the outer shell and an outer circumferential surface of the combustor basket. A plurality of main burners disposed at equal intervals around a central axis of the combustor and configured to inject fuel for premixed combustion, and a pilot burner disposed at a position of a central axis of the combustor are provided in the combustor basket.

In the combustor configured to perform such premixed combustion, combustion oscillations may occur. As a technology of reducing combustion oscillations, for example, Patent Document 1 discloses a combustion oscillation reducing device configured of a hole provided in at least one of the combustor, the duct section connected to the combustor, and the casing section, a filling member disposed to close the hole and having low acoustic impedance and high fluid impedance, and a valve configured to control an aperture of the hole. Acoustic properties (in particular, a frequency) of the combustor are changed when the aperture of the hole is varied. In general, the frequency is increased as the hole is opened. The combustion oscillation reducing device disclosed in Patent Document 1 changes a frequency, a mode, and damping to reduce combustion oscillations by adjusting valves and controlling an aperture of the hole.

In addition, Patent Document 2 discloses a combustion oscillation reducing device (a sound attenuator) configured of an annular pipe body arranged on an outer side of a rear end of a combustor (an outer part of a casing in which the combustor is provided), a throat configured to allow communication between the annular pipe body and the combustor, and a resistor having a plurality of through-holes provided in an end portion of the throat on the side of the combustor. In the combustion oscillation reducing device disclosed in Patent Document 2, fluid particles that are vibration elements of combustion oscillations generated in the combustion region in the combustor resonate with the air in a tubular pipe body connected by the throat, and vibrate in the vicinity of the resistor. Vibrations of the fluid particles in the combustor are attenuated by the vibrations, and combustion oscillations are reduced.

PATENT DOCUMENTS

[Patent Document 1] Japanese Patent No. 3233798

[Patent Document 2] Japanese Patent No. 3999645

SUMMARY OF THE INVENTION

Combustion oscillations in the combustor and an oscillation condition of the pressure fluctuation are expressed by the next Equation (1) of Rayleigh Index R when the pressure fluctuation is expressed by Δp and the heat generation fluctuation is expressed by Δq. Here, T is a period of the fluctuation.

$\begin{array}{cc}R=\frac{1}{T}{\int }_{t-T}^t\Delta p\left(\tau \right)\Delta q\left(\tau \right)dr>0& \left(1\right)\end{array}$

A magnitude of a left side of Equation (1) depends on a pressure mode, a position of a heat source, and a time delay (a natural frequency of an acoustic system and a time delay between a supply system and a combustion system). The combustion oscillation reducing device can adjust the pressure mode and the natural frequency of the acoustic system by adjusting a magnitude, an attachment place, a volume, or the like, of the attachment hole, and can suppress occurrence of the vibrations by reducing a value of the left side of the Equation (1). In addition, since attenuation of a field on the right side of the Equation (1) is increased by opening the hole, it is possible to prevent the combustion oscillations.

Since it will be apparent from Equation (1) that the Rayleigh Index R is a covariance of the pressure and the heat generation, it will be expressed as Equation (2).

$\begin{array}{cc}\begin{array}{c}\frac{1}{T}{\int }_{t-T}^t\Delta p\left(\tau \right)\Delta q\left(\tau \right)d\tau =E\left(\Delta p\Delta q\right)\\ =\mathrm{Cov}\left(p,q\right)\end{array}& \left(2\right)\end{array}$

In order to be stable, Δq when Δp is prior information, i.e., Δq|Δp of Equation (3) must have an opposite sign from Δp.

$\begin{array}{cc}\Delta q|\Delta p=\frac{\mathrm{Cov}\left(p,q\right)}{\mathrm{Var}\left(p\right)}\Delta p& \left(3\right)\end{array}$

That is, when pressure dependency of the heat generation q is expressed as Equation (4), kq<0 is a condition for stability according to the Rayleigh Index R.

Δq=k_qΔ_p  (4)

However, in the Rayleigh Index R, there is no term related to the flow of the air (compressed air), and only the heat generation q is considered. For this reason, a technology of preventing the combustion oscillation more effectively with taking into consideration of the fluctuation of the flow of the air in the combustion oscillations has been required.

In consideration of the above-mentioned problems, the present disclosure is directed to providing a gas turbine capable of more effectively suppressing combustion oscillations.

According to an aspect of the present disclosure, a gas turbine includes a rotor that is rotatable about an axis; a casing configured to cover the rotor in a circumferential direction and having an annular space therein; a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing; a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas; a turbine driven by the combustion gas; and an air introduction passage defined by a partition plate configured to divide the space in the casing in the circumferential direction of the rotor and an inner circumferential surface of the casing and configured to introduce the compressed air in the casing into the combustor.

According to an aspect of the present disclosure, a gas turbine includes a rotor that is rotatable about an axis; a casing configured to cover the rotor in a circumferential direction and having an annular space therein; a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing; a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas; a turbine driven by the combustion gas; and an annular air introduction passage defined by a first cylinder section that surrounds the combustor and a second cylinder section that surrounds the first cylinder section and configured to introduce the compressed air in the casing into the combustor.

According to an aspect of the present disclosure, a gas turbine includes a rotor that is rotatable about an axis; a casing configured to cover the rotor in a circumferential direction and having an annular space therein; a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing; a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas; a turbine driven by the combustion gas; and an air introduction passage configured to introduce the compressed air in the casing into the combustor, the air introduction passage being defined by a first wall section having an L-shaped cross section configured of a first plate section extending from an outer circumferential surface of the combustor and a second plate section extending from the first plate section toward a downstream side of the rotor in the axial direction, and a second wall section having a U-shaped cross section configured of a third plate section extending from an inner circumferential surface of the casing toward the downstream side of the rotor in the axial direction, a fourth plate section extending from the third plate section toward the outer circumferential surface of the combustor and a fifth plate section extending from the fourth plate section toward the upstream side of the rotor in the axial direction, and the second plate section of the first wall section being disposed between the third plate section and the fifth plate section of the second wall section.

According to an aspect of the present disclosure, a gas turbine includes a rotor that is rotatable about an axis; a casing configured to cover the rotor in a circumferential direction and having an annular space therein; a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing; a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas; a turbine driven by the combustion gas; and an air introduction passage defined by a plurality of guide pipes having a first opening section connected to an inlet port of the combustor and a second opening section that opens in the space in the casing on the downstream side in the axial direction of the rotor.

According to the gas turbine of the present disclosure, combustion oscillations can be more effectively suppressed.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic diagram showing a configuration of a gas turbine according to a first embodiment of the present disclosure.

FIG. 2 is a cross-sectional view showing a schematic configuration around a combustor of the gas turbine according to the first embodiment of the present disclosure along a rotor axis.

FIG. 3 is a view showing a model of the combustor according to the first embodiment of the present disclosure.

FIG. 4 is a view showing an example of a stability condition of the combustor according to the first embodiment of the present disclosure.

FIG. 5 is a cross-sectional view showing a schematic configuration around the combustor of the gas turbine according to the first embodiment of the present disclosure when seen in a rotor axis direction.

FIG. 6 is a schematic diagram around the combustor of the gas turbine according to the first embodiment of the present disclosure.

FIG. 7 is a schematic diagram around a combustor of a gas turbine according to Variant 1 of the first embodiment of the present disclosure.

FIG. 8 is a schematic diagram around a combustor of a gas turbine according to Variant 2 of the first embodiment of the present disclosure.

FIG. 9 is a cross-sectional view showing a schematic configuration around a combustor of a gas turbine according to Variant 3 of the first embodiment of the present disclosure when seen in the rotor axis direction.

FIG. 10 is a schematic diagram around the combustor of the gas turbine according to Variant 3 of the first embodiment of the present disclosure.

FIG. 11 is a schematic diagram around a combustor of a gas turbine according to Variant 4 of the first embodiment of the present disclosure.

FIG. 12 is a cross-sectional view showing a schematic configuration around a combustor of a gas turbine according to a second embodiment of the present disclosure along a rotor axis.

FIG. 13 is a cross-sectional view showing a schematic configuration around a combustor of a gas turbine according to a third embodiment of the present disclosure along a rotor axis.

FIG. 14 is a cross-sectional view showing a schematic configuration around a combustor of a gas turbine according to a fourth embodiment of the present disclosure along a rotor axis.

FIG. 15 is a cross-sectional view showing a schematic configuration of a combustor according to a fifth embodiment of the present disclosure along a combustor axis.

FIG. 16 is a cross-sectional view showing a schematic configuration of the combustor according to the fifth embodiment of the present disclosure when seen in a combustor axis direction.

FIG. 17 is a cross-sectional view showing a schematic configuration of a combustor according to a variant of the fifth embodiment of the present disclosure along a combustor axis.

FIG. 18 is a cross-sectional view showing a schematic configuration of a combustor according to a sixth embodiment of the present disclosure along a combustor axis.

FIG. 19 is a cross-sectional view showing a schematic configuration of a combustor according to a seventh embodiment of the present disclosure along a combustor axis.

FIG. 20 is a cross-sectional view showing a schematic configuration of the combustor according to the seventh embodiment of the present disclosure when seen in a combustor axis direction.

DETAILED DESCRIPTION OF THE INVENTION

First Embodiment

Hereinafter, a gas turbine according to a first embodiment of the present disclosure will be described with reference to the accompanying drawings.

(Entire Configuration)

FIG. 1 is a schematic diagram showing a configuration of the gas turbine of the first embodiment of the present disclosure.

As shown in FIG. 1, a gas turbine 1 according to the embodiment includes a rotor 4, a casing 5, a compressor 2, a plurality of combustors 10, and a turbine 3.

The rotor 4 rotates about a rotor axis Ar, and connects the compressor 2 and the turbine 3. Here, a direction in which the rotor axis Ar extends is referred to as a rotor axis direction Da, one side of both ends in the rotor axis direction Da is referred to as an axis upstream side Dau, and the other side is referred to as an axis downstream side Dad. A circumferential direction around the rotor axis Ar is referred to as a rotor circumferential direction Dcr. In addition, a direction perpendicular to the rotor axis Ar is referred to as a rotor radial direction Drr. In the rotor radial direction Drr, a side close to the rotor axis Ar is referred to as a rotor inner side Drri, and a side away from the rotor axis Ar is referred to as a rotor outer side Drro.

The casing 5 covers the rotor 4 in a circumferential direction Dcr, and has an annular space therein.

The compressor 2 takes in external air as a working fluid, generates compressed air (air) A, and sends the air into the casing 5.

As shown in FIG. 1, the plurality of combustors 10 are disposed in the casing 5 at equal intervals in the circumferential direction of the rotor 4. Each of the combustors 10 is in communication with an outlet port of the compressor 2, and generates a high temperature and a high pressure of combustion gas G by mixing and combusting fuel and the compressed air A supplied from the compressor 2.

The turbine 3 converts thermal energy of the combustion gas G delivered from the combustors 10 into rotation energy of the rotor 4. Then, the rotation energy is transmitted to a generator (not shown) connected to the rotor 4.

Further, in the combustors 10, the combustor axes Ac are disposed radially in a state in which an inlet port-side of the combustor 10 is inclined with respect to a rotation axis Ar of the rotor 4 in the gas turbine 1 away from the outlet port-side in the radial direction.

FIG. 2 is a cross-sectional view showing a schematic configuration around the combustor of the gas turbine according to the first embodiment of the present disclosure along a rotor axis.

As shown in FIG. 2, each of the combustors 10 includes an outer shell 11, a combustor basket 12, a plurality of burners 13, and a transition piece 15.

The outer shell 11, the combustor basket 12, and the transition piece 15 are disposed in a space in the casing 5 that is a dual annular closed space coaxial with the rotor 4. In addition, the outer shell 11, the combustor basket 12, and the transition piece 15 are all cylindrical around the combustor axis Ac.

Here, a direction in which the combustor axis Ac extends is referred to as an axis direction Dc, one side of both ends in the axis direction Dc is referred to as a tip side Dct, and the other side is referred to as a base side Dcb. As shown in FIG. 1 and FIG. 2, the tip side Dct is the axis downstream side Dad in the rotor axis direction Da, and the base side Dcb is the axis upstream side Dau in the rotor axis direction Da. Further, the combustor axis Ac is inclined with respect to the rotor axis Ar to approach the rotor axis Ar as it goes toward the tip side Dct. In addition, a circumferential direction around the combustor axis Ac is referred to as a circumferential direction Dcc. In addition, a direction perpendicular to the combustor axis Ac is referred to as a radial direction Drc. A side close to the combustor axis Ac is an inner circumferential side or a radial direction inner side Drci, and a side away from the combustor axis Ac is referred to as an outer circumferential side or a radial direction outer side Drco.

The outer shell 11 has a flange 11f extending from the combustor axis Ac to the radial direction outer side Drco. The flange 11f is attached to the casing 5 by bolts to close a combustor attachment hole provided in the casing 5.

The combustor basket 12 is attached to an inner circumferential side of the outer shell 11 with a gap from the flange 11f. The plurality of burners 13 are disposed on an inner circumferential side of the combustor basket 12. The transition piece 15 is connected to the tip side Dct of the combustor basket 12.

The plurality of burners 13 (13a, 13b) extend in the axis direction Dc, and holes configured to inject fuel are formed therein. The plurality of burners 13 are all fixed to the outer shell 11. The plurality of burners 13 are configured of pilot burners 13a and main burners 13b. The pilot burners 13a are disposed on the combustor axis Ac. The plurality of main burners 13b are arranged around the pilot burners 13a at equal intervals in the circumferential direction Dcc.

(With Respect to Combustion Oscillations)

FIG. 3 is a view showing a model of the combustor according to the first embodiment of the present disclosure.

The model shown in FIG. 3 represents the combustor 10 by explicitly centralizing a flow of the air (the compressed air A). In FIG. 3, a point A shows an inlet port-side boundary point of the combustor 10 (a boundary point on the side of the compressor 2), a point B shows a point representing a spatial extent of the combustors 10, and a point C shows an outlet port-side boundary point of the combustor 10 (a boundary point on the side of the turbine 3). At the point B, there are air g_A (kg/s) inflowing from the inlet port-side boundary point A and air g_B (kg/s) outflowing to the outlet port-side boundary point B. Specific enthalpies of air of the inflow and the outflow are h_A (J/kg) and h_B (J/kg). A volume of the combustors 10 is V (m³), and a mass of air of the combustors 10 is m_B (kg).

A mass balance of the combustors 10 is expressed by the following Equation (5). ⋅ on the reference sign represents time differentiation. For example, m_BÂ⋅ is a time rate of change (kg/s) of the mass m_B (kg).

{right arrow over (m)}_B=g_A−g_B  (5)

An energy balance of the combustor is expressed by the following Equation (6). In order to simplify the description, the fuel is considered for the heat generation q only, and the mass of the fuel is not considered.

m_B{dot over (h)}_B=(h_Ag_A−h_Bg_B+q+{dot over (p)}_BV)−{dot over (m)}_Bh_B  (6)

Since a specific volume of the air in the combustor is expressed as v_B (p_B, h_B), the pressure and the specific enthalpy have a relationship expressed by the next Equation (7).

$\begin{array}{c}\dot{V}={\dot{m}}_B{v}_B+m_B{\stackrel{.}{v}}_B\\ {\stackrel{.}{m}}_B{v}_B+{m_B\left(\frac{\partial v}{\partial p}\right)}_h{\dot{p}}_B+{m_B\left(\frac{\partial v}{\partial h}\right)}_p{\dot{h}}_B\end{array}$

Since a volume V does not change over time, a time differential value of V is zero. Accordingly, the following Equation (8) for the pressure of the combustor is obtained.

$\begin{array}{cc}\begin{array}{c}0={\dot{m}}_B{v}_B+{m_B\left(\frac{\partial v}{\partial p}\right)}_h{\stackrel{.}{p}}_B+{m_B\left(\frac{\partial v}{\partial h}\right)}_p{\stackrel{.}{h}}_B\\ =\left(g_A-g_B\right)v_B+{m_B\left(\frac{\partial v}{\partial p}\right)}_h{\stackrel{.}{p}}_B+{\left(\frac{\partial v}{\partial h}\right)}_p\left(h_A{g}_A-h_B{g}_B+q+{\stackrel{.}{p}}_{B}V-{\dot{m}}_B{h}_B\right)\end{array}& \left(8\right)\end{array}$

Here, a setting state of the combustor is expressed by a flow rate g_{A, e}, a specific enthalpy h_{A, e}, and heat generation qe of the inflow air, and the fluctuation of the state of the combustors 10 is expressed by the pressure fluctuation Δp_B of the combustor as shown in the following Equation (9).

g_A=k_gAΔp_B+g_A,e

g_Bk_gBΔp_Bg_A,e

h_B=k_hBΔp_B+h_A,e+q_eg_A,e⁻¹

q=k_qΔp_Bq_e  (9)

When this is substituted into the equation for the pressure, a linear approximation model of the following Equation (10) is obtained with respect to dynamic characteristics of the pressure.

$\begin{array}{cc}\begin{array}{c}0={\stackrel{.}{m}}_B{v}_B+{m_B\left(\frac{\partial v}{\partial p}\right)}_h{\stackrel{.}{p}}_B+{m_B\left(\frac{\partial v}{\partial h}\right)}_p{\stackrel{.}{h}}_B\\ =\begin{array}{c}\left(k_{\mathrm{gA}}-k_{\mathrm{gB}}\right)v_{B,e}\Delta p_B+{m_{B,e}\left(\frac{\partial v}{\partial p}\right)}_h\Delta {\stackrel{.}{p}}_B+{V\left(\frac{\partial v}{\partial p}\right)}_h\Delta {\stackrel{.}{p}}_B+\\ {\left(\frac{\partial v}{\partial h}\right)}_p\left(h_{A,e}{k}_{\mathrm{gA}}-h_{B,e}{k}_{\mathrm{gB}}+k_q-\left(k_{\mathrm{gA}}-k_{\mathrm{gB}}\right)h_{B,e}\right)\Delta p_B\end{array}\\ =\begin{array}{c}\left(k_{\mathrm{gA}}-k_{\mathrm{gB}}\right)v_{B,e}\Delta p_B+\left({m_{B,e}\left(\frac{\partial v}{\partial p}\right)}_h+{V\left(\frac{\partial v}{\partial h}\right)}_p\right)\Delta {\stackrel{.}{p}}_B+\\ {\left(\frac{\partial v}{\partial h}\right)}_p\left(\left(h_{A,e}-h_{B,e}\right)k_{\mathrm{gA}}+k_q\right)\Delta p_B\end{array}\\ =\begin{array}{c}\left(k_{\mathrm{gA}}-k_{\mathrm{gB}}\right)v_{B,e}\Delta p_B+\left({m_{B,e}\left(\frac{\partial v}{\partial p}\right)}_h+{V\left(\frac{\partial v}{\partial h}\right)}_p\right)\Delta {\stackrel{.}{p}}_B+\\ {\left(\frac{\partial v}{\partial h}\right)}_p\left(-q_e{g}_{A,e}^{-1}{k}_{\mathrm{gA}}+k_q\right)\Delta p_B\end{array}\\ =\begin{array}{c}\left(\left(v_{B,e}-{\left(\frac{\partial v}{\partial h}\right)}_p{q}_e{g}_{A,e}^{-1}\right)k_{\mathrm{gA}}-v_{B,e}{k}_{\mathrm{gB}}+{\left(\frac{\partial v}{\partial h}\right)}_p{k}_q\right)\Delta p_B+\\ \left({m_{B,e}\left(\frac{\partial v}{\partial p}\right)}_h+{V\left(\frac{\partial v}{\partial h}\right)}_p\right)\Delta {\stackrel{.}{p}}_B\end{array}\end{array}& \left(10\right)\end{array}$

When the solution of the linear approximation model is represented as Equation (11) and substituted into the above-mentioned Equation (10), the characteristic equation represented by Equation (12) is obtained.

Δp_B(t)=e^λt  (11)

$\begin{array}{cc}0=\left(\left(v_{B,e}-{\left(\frac{\partial v}{\partial h}\right)}_p{q}_e{g}_{A,e}^{-1}\right)k_{\mathrm{gA}}-v_{B,e}{k}_{\mathrm{gB}}+{\left(\frac{\partial v}{\partial h}\right)}_p{k}_q\right)+\left({m_{B,e}\left(\frac{\partial v}{\partial p}\right)}_h+{V\left(\frac{\partial v}{\partial h}\right)}_p\right)\lambda & \left(12\right)\end{array}$

Then, a characteristic root λ is expressed as Equation (13).

$\begin{array}{cc}\lambda =\frac{\left(v_{B,e}-{\left(\frac{\partial v}{\partial h}\right)}_p{q}_e{g}_{A,e}^{-1}\right)k_{\mathrm{gA}}-v_{B,e}{k}_{\mathrm{gB}}+{\left(\frac{\partial v}{\partial h}\right)}_p{k}_q}{-{m_{B,e}\left(\frac{\partial v}{\partial p}\right)}_h-{V\left(\frac{\partial v}{\partial h}\right)}_p}& \left(13\right)\end{array}$

The dynamic characteristics of the pressure are unstable when a characteristic root is positive, and they are stable when it is negative. Positive or negative of the characteristic root will be described separately for each case.

Case 1 is a case of modeling assuming that the flow rate of the air is constant. If the flow rate is constant, k_gA and k_gB are zero in the above-mentioned Equation (13). In this case, the characteristic root is the following Equation (14).

$\begin{array}{cc}\begin{array}{cc}\lambda =\frac{{\left(\frac{\partial v}{\partial h}\right)}_p{k}_q}{-{m_{B,e}\left(\frac{\partial v}{\partial p}\right)}_h-{V\left(\frac{\partial v}{\partial h}\right)}_p}& \left(\mathrm{when}{k}_{\mathrm{gA}}=k_{\mathrm{gB}}=0\right)\end{array}& \left(14\right)\end{array}$

(dv/dp)_h is negative since the gas contracts when the pressure is increased, and (dv/dh)_p is positive since the gas expands when the gas is heated. However, since a term of (dv/dp)_h in a denominator, which is a load sum of both, is dominated, a sign of the denominator is positive finally. Accordingly, the condition in which the characteristic root is negative, i.e., a stability condition is expressed by Equation (15), and has the same result as the Rayleigh Index described above.

k_q<0  (15)

From Case 1, it is inferred that the Rayleigh Index expresses the stability condition when the flow rate of the air does not change and only the heat generation changes.

Case 2 is a more realistic case when the flow rate of the air changes depending on the pressure of the combustors 10. In the embodiment, it is focused on Case 2. In Case 2, the condition for stability is that a numerator (i.e., a load sum of k_q, k_gA and k_gB) of Equation (16) of the characteristic root is positive.

$\begin{array}{cc}\begin{array}{cc}\lambda =\frac{\begin{array}{c}\left(v_{B,e}-{\left(\frac{\partial v}{\partial h}\right)}_p{q}_e{g}_{A,e}^{-1}\right)k_{\mathrm{gA}}-\\ v_{B,e}{k}_{\mathrm{gB}}+{\left(\frac{\partial v}{\partial h}\right)}_p{k}_q\end{array}}{-{m_{B,e}\left(\frac{\partial v}{\partial p}\right)}_h-{V\left(\frac{\partial v}{\partial h}\right)}_p}& \left(\mathrm{when}{k}_{\mathrm{gA}},k_{\mathrm{gB}}\mathrm{is}\mathrm{not}\mathrm{zero}\right)\end{array}& \left(16\right)\end{array}$

A load coefficient of k_gB is (−v_{B, e}), and since the value is negative, an effect of preventing the combustion oscillations is increased as k_gB is increased. That is, it is better if the air flows easily to the outlet port-side. The load coefficient of k_gA is (v_{B, e}−(dv/dh)_pq_e/g_{A, e}). In the gas turbine in recent years, since the outlet port temperature of the combustor is set to a high temperature for high efficiency and the absolute temperature with respect to the combustor outlet/inlet port is about three times the value, the load coefficient of k_gA is about (⅓)×v_{B, e}. That is, in the gas turbine in recent years in which the absolute temperature TB of the outlet port of the combustor is about three times the absolute temperature TA of the inlet port, the characteristic equation is expressed as Equation (17). It is important that positive or negative of the characteristic root also depends on k_gA and k_gB, which represent the flow rate of the air.

$\begin{array}{cc}\begin{array}{cc}\lambda =\frac{\begin{array}{c}{v}_{B,e}\left(\frac{1}{3}{k}_{\mathrm{gA}}-k_{\mathrm{gB}}\right)+\\ {\left(\frac{\partial v}{\partial h}\right)}_p{k}_q\end{array}}{\begin{array}{c}-m_{B,e}{\left(\frac{\partial v}{\partial p}\right)}_h-\\ V{\left(\frac{\partial v}{\partial h}\right)}_p\end{array}}& \left(\mathrm{when}{k}_{\mathrm{gA}},k_{\mathrm{gB}}\mathrm{is}\mathrm{not}\mathrm{zero}\mathrm{and}{T}_B\approx 3T_A\right)\end{array}& \left(17\right)\end{array}$

As expected at the beginning, when modeling realistically, it turned out that the flow rate of the air is also related to the combustion oscillations. Based on this result, avoidance of the combustion oscillations is examined.

Conservation of the momentum of the inflow air g_A and the outflow air g_B is expressed by the following Equation (18). Friction of the air is ignored for simplicity.

$\begin{array}{cc}\begin{array}{c}{\int }_0^{L_A}{A}_A^{-1}dL·{\stackrel{.}{g}}_A=p_A-p_B\\ {\int }_0^{L_B}{A}_B^{-1}dL·{\stackrel{.}{g}}_B=p_B-p_C\end{array}& \left(18\right)\end{array}$

In order to simplify the description, the air is divided into a base line of a flow rate in which g_A and g_B are simultaneously increased and decreased and a relative flow rate in which both are differently increased and decreased in a seesaw shape such that movement of the point mass system is divided into movement of a center of gravity and relative movement. The g_A and g_B are expressed as the following Equation (19) by setting the flow rate of the base line to g_{A, e} and the relative flow rate to ξ_A and ξ_B.

$\begin{array}{cc}\{\begin{array}{c}{g}_A=g_{A,e}+{\xi }_A\\ g_B=g_{A,e}+{\xi }_B\end{array}& \left(19\right)\end{array}$

A total momentum of the flow rate of the base line is expressed by the following Equation (20) by substituting g_{A, e} into the equation of the momentum of the g_A and g_B and summing them. Since the flow rate of the base line does not depend on the pressure p_B of the combustor, the flow rate of the base line is independent of the combustion oscillations. An integrally calculated portion of the following Equation (20) expresses an equivalent length of the flow channel with respect to inertia.

$\begin{array}{cc}\left({\int }_0^{L_A}{A}_A^{-1}dL+{\int }_0^{L_B}{A}_B^{-1}dL\right){\stackrel{.}{g}}_{A,e}=p_A-p_C& \left(20\right)\end{array}$

Meanwhile, since the total momentum of the relative movement of the g_A and g_B caused by the combustion oscillations is expressed by a difference between the momentum of ξ_A and the momentum of ξ_B as shown in the following Equation (21) because they move in opposite directions. Since the flow rate by the relative movement reflects compression and expansion of the gas for the combustor, it represents the combustion oscillations.

$\begin{array}{cc}{\stackrel{.}{\xi }}_A{\int }_0^{L_A}{A}_A^{-1}dL-{\stackrel{.}{\xi }}_B{\int }_0^{L_B}{A}_B^{-1}dL=p_A-2p_B-p_C& \left(21\right)\end{array}$

In the combustion oscillation, since only the two flow channels, inflow and outflow, of the combustor exchange the momentum locally, the sum of the momentums of both is zero as shown in Equation (22).

$\begin{array}{cc}{\stackrel{.}{\xi }}_A{\int }_0^{L_A}{A}_A^{-1}dL+{\stackrel{.}{\xi }}_B{\int }_0^{L_B}{A}_B^{-1}dL=0& \left(22\right)\end{array}$

For simplicity of notation, like Equation (23), a ratio of the equivalent lengths of the flow channels of the inflow and the outflow is expressed as R_AB.

$\begin{array}{cc}{R}_{\mathrm{AB}}=\frac{{\int }_0^{L_A}{A}_A^{-1}dL}{{\int }_0^{L_B}{A}_B^{-1}dL}& \left(23\right)\end{array}$

Since the sum of the momentums of both the inflow and the outflow becomes zero, a flow rate due to relative movement is expressed as the following Equation (24).

{dot over (ξ)}_B=−R_AB{dot over (ξ)}_A  (24)

Since a ratio of ξ_A and ξ_B is equal to a ratio of k_gB and k_gA, the following Equation (25) is obtained.

k_gB=−R_ABk_gA  (25)

When it is substituted into Equation (17), the characteristic root is expressed as Equation (26).

$\begin{array}{cc}\begin{array}{cc}\lambda =\frac{\begin{array}{c}{v}_{B,e}\left(\frac{1}{3}+R_{\mathrm{AB}}\right)k_{\mathrm{gA}}+\\ {\left(\frac{\partial v}{\partial h}\right)}_p{k}_q\end{array}}{\begin{array}{c}-m_{B,e}{\left(\frac{\partial v}{\partial p}\right)}_h-\\ V{\left(\frac{\partial v}{\partial h}\right)}_p\end{array}}& \left(\mathrm{when}{k}_{\mathrm{gA}},k_{\mathrm{gB}}\mathrm{is}\mathrm{not}\mathrm{zero}\mathrm{and}{T}_B\approx 3T_A\right)\end{array}& \left(26\right)\end{array}$

Since the value of the denominator of Equation (26) is positive, the stability condition is given by the following Equation (27).

$\begin{array}{cc}{k}_q<-v_{B,e}\left(\frac{1}{3}+R_{\mathrm{AB}}\right)k_{\mathrm{gA}}/{\left(\frac{\partial v}{\partial h}\right)}_p& \left(27\right)\end{array}$

When a fluid of the combustor is approximated by an ideal gas, the above-mentioned Equation (27) is expressed as the following Equation (28).

$\begin{array}{cc}{k}_q<-h_{B,e}\left(\frac{1}{3}+R_{\mathrm{AB}}\right)k_{\mathrm{gA}}& \left(28\right)\end{array}$

From the Equation (28), it can be seen that the stability depends on the ratio R_AB of the equivalent length of the flow channels of the inflow and the outflow.

FIG. 4 is a view showing an example of a stability condition of the combustor according to the first embodiment of the present disclosure.

Since a value of a flow rate coefficient k_gA of the inflow is negative, a stability condition as shown in FIG. 4 is obtained for a region of k_gA<0. A broken line L1 of FIG. 4 shows a range of stability when R_AB=0. A solid line L2 shows a range of stability when R_AB=1. It can be seen that the range of stability expands when the ratio R_AB of the equivalent length of the flow channel of the inflow and the outflow increases. Based on such findings, in the embodiment, the gas turbine 1 including the configuration of avoiding the combustion oscillation will be described from the viewpoint of the ratio R_AB of the equivalent length of the flow channel.

(With Respect to Flow of Compressed Air)

FIG. 2 shows an aspect in which the compressed air A in the space in the casing 5 flows through a meridian plane including the rotor axis Ar and the combustor axis Ac. In the casing 5, a compressor outlet port flow channel F1 is connected to the inner side Drri in the radial direction of the rotor 4. The compressed air A compressed by the compressor 2 flows into the casing 5 from the compressor outlet port flow channel F1. The compressed inflow air A reverses a flow direction of the rotor axis direction Da and is directed toward an inlet port of the combustor 10 (the base side Dcb). An annular space between an inner circumferential surface 11a of the outer shell 11 and an outer circumferential surface 12a of the combustor basket 12 forms a part of an air introduction passage F2 configured to introduce the compressed air A in the casing 5 into the combustor basket 12. In addition, the combustor basket 12 is disposed with a gap between the combustor basket 12 and the flange 11f of the outer shell 11. The compressed air A in the air introduction passage F2 flows into the combustor basket 12 from the gap between the combustor basket 12 and the flange 11f. The compressed inflow air A in the combustor basket 12 flows out into the transition piece 15. Fuel is injected into the transition piece 15 from the burners 13. The fuel is mixed with the compressed air A to be combusted in the transition piece 15 to generate the combustion gas G. The combustion gas G is introduced into the turbine 3 by the transition piece 15.

In a conventional gas turbine, the space in the casing is common to the plurality of combustors and there are no partitions inside. For this reason, the compressed air flows freely in the casing in the circumferential direction of the rotor, i.e., also in a direction perpendicular to the meridian plane.

When there are the plurality of combustors, it is inevitable that the pressure in the combustion chamber of each combustor will differ. When the compressed air is distributed in the combustors from the space in the casing common to the plurality of combustors (hereinafter, also referred to as “an air introduction passage”), if the internal pressure in some of the combustors is increased, supply of the compressed air to the corresponding combustor is reduced, and the decrement bypasses the other combustors and balances them.

As described in the above-mentioned Equation (26), considering that the combustion oscillations easily occur as the intake airflow rate is decreased when the combustor pressure is increased, it is not preferable that the internal pressure is increased in one of the combustors and the compressed air supplied thereinto bypasses to the other combustor, from the viewpoint of the combustion oscillations.

(With Respect to Configuration of Suppressing Combustion Oscillations)

In the related art, in order to suppress the combustion oscillations, an acoustic attenuator is attached to the combustor main body. On the other hand, in the embodiment, based on the above-mentioned knowledge, it is focused on the air introduction passage to the combustors 10. The focus is placed on the result of the study of the case 2 based on the Rayleigh Index, which shows that the combustion oscillations tend to occur as the intake airflow rate decreases when the combustor pressure rises. That is, the gas turbine 1 according to the embodiment suppresses reduction in intake airflow rate of the combustors 10 even when the pressure rises in one of the plurality of combustors 10. Alternatively, when the pressure in the combustor is decreased, an increase in intake airflow rate of the combustor 10 is suppressed. For this purpose, in the embodiment, the equivalent length for the inertia of the air introduction passage to the combustors 10 is increased. Further, while the differential pressure of the air introduction passage may be increased in addition to the increase in equivalent length for the inertia, an increase in differential pressure may reduce the efficiency of the heat engine.

FIG. 5 is a cross-sectional view showing a schematic configuration around the combustor of the gas turbine according to the first embodiment of the present disclosure when seen in the rotor axis direction.

FIG. 6 is a schematic diagram around the combustor of the gas turbine according to the first embodiment of the present disclosure.

As shown in FIG. 5 and FIG. 6, the gas turbine 1 according to the embodiment further includes a partition plate 20 configured to divide a space in the casing 5 in the circumferential direction Dcr of the rotor 4, and an air introduction passage F2 defined by inner circumferential surfaces 5a and 5b of the casing 5 and configured to introduce the compressed air A in the casing 5 into the combustors 10.

The partition plate 20 extends in the rotor radial direction Drr from the inner circumferential surface 5a of the casing 5 on an outer side Drro in the rotor radial direction to the inner circumferential surface 5b on the inner side Drri in the rotor radial direction. In addition, the partition plate 20 extends from an inlet port (the base side Dcb) of the combustor basket 12 to an outlet port (the tip side Dct) of the transition piece in the rotor axis direction Da.

The air introduction passage F2 circulates the compressed inflow air A to the base side Dcb of the combustors 10 from the compressor outlet port flow channel F1 and introduces the compressed inflow air A into the combustors 10 (the combustor basket 12).

Further, in the embodiment, the partition plate 20 defines the space in the casing 5 to the air introduction passages F2 independent for each of the combustors 10 by providing them in intermediate portions between the plurality of combustors 10. In this way, if the air introduction passage F2 is independently formed for each of the combustors 10, when an internal pressure of a certain combustor 10 is increased, it is able to prevent the compressed air A from bypassing this combustor 10 and flowing into the other combustors 10. Accordingly, the combustion oscillations can be suppressed.

In addition, in the related art, when the internal pressure of the combustor is increased and a supply flow rate of the compressed air into the combustor is decreased, the compressed air may cause abnormal combustion, referred to as backfire in which flame runs up. However, the gas turbine 1 according to the embodiment reduces the bypass of the compressed air A and suppresses a decrease in supply flow rate to the combustors 10 even when an internal pressure in a certain combustor 10 is increased. Accordingly, the gas turbine 1 according to the embodiment is capable of suppressing occurrence of backfire in the combustors 10.

Further, while FIG. 5 and FIG. 6 show an example in which the partition plate 20 is disposed not to interfere with the combustors 10, there is no limitation thereto. In another embodiment, the partition plate 20 may be inclined with respect to the meridian plane and arranged across the combustors 10.

In addition, while FIG. 5 and FIG. 6 show an example in which the partition plates 20 are provided between the neighboring combustors 10 and the air introduction passage F2 is formed independently for each of the combustors 10, there is no limitation thereto. In the other embodiment, the partition plate 20 may be reduced. For example, the neighboring two combustors 10 may be provided as one set, the partition plate 20 may be provided between the sets, and the air introduction passage F2 may be formed independently for each set. Accordingly, the air introduction passage F2 can be formed with a simpler configuration.

(Effects)

As described above, the gas turbine 1 according to the embodiment includes the rotor 4 that is rotatable about an axis Ar, the casing 5 configured to cover the rotor 4 in the circumferential direction Dcr and having an annular space therein, the compressor 2 configured to generate a high pressure of the compressed air A obtained by compressing external air and send the compressed air A into the casing 5, the plurality of combustors 10 disposed in the casing 5 at equal intervals in the circumferential direction Dcr of the rotor 4 and configured to combust the compressed air A and the fuel take in from the casing 5 and generate the combustion gas G, the turbine 3 driven by the combustion gas G, the partition plate 20 configured to divide the space in the casing 5 in the circumferential direction Dcr of the rotor 4, and the air introduction passage F2 defined by the inner circumferential surface 5a of the casing 5 and configured to introduce the compressed air A in the casing 5 into the combustors 10.

By providing such a configuration, when an internal pressure of a certain combustor 10 among the plurality of combustors is increased, the gas turbine 1 can suppress the compressed air A in the casing 5 from bypassing from this combustor 10 to the other combustors 10. Accordingly, it is possible to suppress a decrease and an increase in supply flow rate of the compressed air A to each of the combustors 10 due to fluctuation of the combustor pressure. In addition, since it is possible to suppress reduction of the compressed air A supplied to the combustors 10 in which the internal pressure is increased, it is possible to suppress combustion oscillations in the combustors 10 and occurrence of the backfire.

In addition, the partition plate 20 is provided in the middle of each of the plurality of combustors.

Accordingly, since the gas turbine 1 can have the air introduction passage F2 independent for each of the combustors 10, even when the internal pressure of each of the combustors 10 differs, it is possible to suppress occurrence of a difference in flow rate of the compressed air A supplied to each of the combustors 10. Accordingly, it is possible to more reliably suppress the combustion oscillations.

<Variant 1 of First Embodiment>

Next, the gas turbine 1 according to Variant 1 of the first embodiment will be described.

The same components as the above-mentioned embodiment are designated by the same reference signs and detailed description thereof will be omitted.

FIG. 7 is a schematic diagram around a combustor of the gas turbine according to Variant 1 of the first embodiment of the present disclosure.

As shown in FIG. 7, in the gas turbine 1 according to the embodiment, a baffle plate 21 configured to bypass the compressed air A is provided in the air introduction passage F2.

At least one baffle plate 21 is provided in the air introduction passage F2. The number of baffle plates 21 may be increased or decreased arbitrarily. In the example of FIG. 7, three of the baffle plates 21a, 21b and 21c are provided in sequence from an upstream side (i.e., the axis downstream side Dad) of the air introduction passage F2.

The baffle plates 21a and 21c extend from the inner circumferential surface 5b of the casing 5 on the inner side Drri in the rotor radial direction toward the outer side Drro in the rotor radial direction, and have a gap with the inner circumferential surface 5a of the casing 5 on the outer side Drro in the rotor radial direction.

The baffle plate 21b extends from the inner circumferential surface 5a of the casing 5 on the outer side Drro in the rotor radial direction toward the inner side Drri in the rotor radial direction, and has a gap with the inner circumferential surface 5b of the casing 5 on the inner side Drri in the rotor radial direction.

In addition, the combustor basket 12 and the transition piece 15 of the combustor 10 are disposed to pass through the baffle plates 21a to 21c.

The compressed air A flowing into the casing 5 from the compressor outlet port flow channel F1 collides with the baffle plate 21a and then flows toward the outer side Drro in the rotor radial direction along the baffle plate 21a. In addition, the compressed air A passing through the gap between the baffle plate 21a and the inner circumferential surface 5a collides with the baffle plate 21b and then flows toward the inner side Drri in the rotor radial direction along the baffle plate 21b. Further, the compressed air A passing through the gap between the baffle plate 21b and the inner circumferential surface 5b collides with the baffle plate 21c and then flows toward the outer side Drro in the rotor radial direction along the baffle plate 21c. The compressed air A passing through the gap between the baffle plate 21c and the inner circumferential surface 5a flows into the combustors 10 (the combustor basket 12) through the gap between the inner circumferential surface 11a of the outer shell 11 and the outer circumferential surface 12a of the combustor basket 12. In this way, the compressed air A flowing through the air introduction passage F2 is introduced into the combustors 10 while changing a direction using the baffle plate 21. That is, a flow channel length of the air introduction passage F2 is extended by the baffle plate 21.

In this way, since the baffle plate 21 configured to bypass the compressed air A is provided in the air introduction passage F2, it is possible to lengthen the flow channel length of the air introduction passage F2 (lengthen the equivalent length for inertia). Accordingly, it is possible to suppress a decrease in supply flow rate of the compressed air A due to the increase in internal pressure of the combustor 10 and suppress the combustion oscillations.

<Variant 2 of First Embodiment>

Next, the gas turbine 1 according to Variant 2 of the first embodiment of the present disclosure will be described.

The same components as the above-mentioned embodiment are designated by the same reference signs and detailed description thereof will be omitted.

FIG. 8 is a schematic diagram around a combustor of the gas turbine according to Variant 2 of the first embodiment of the present disclosure.

As shown in FIG. 8, in the gas turbine 1 according to the embodiment, a communication hole 22 in communication with the adjacent air introduction passages F2 in the rotor circumferential direction Dcr is provided in the partition plate 20. For example, a material having flow resistance such as a perforated metal is attached to the partition plate 20 in the communication hole 22.

In this way, since the communication hole 22 is provided in the partition plate 20, when a pressure of a certain combustor 10 is higher than that of the other combustors 10, a flow due to the bypass of the compressed air A is generated in the communication hole 22.

In the related art, since there is no flow resistance in the casing, even when the compressed air bypasses the adjacent combustors, no attenuation effect can be obtained. On the other hand, the gas turbine 1 according to the variant can suppress pressure fluctuations because a bypass flow of the compressed air A is applied as an attenuation force by adding the flow resistance to the communication hole 22.

<Variant 3 of First Embodiment>

Next, the gas turbine 1 according to Variant 3 of the first embodiment will be described.

The same components as the above-mentioned embodiment are designated by the same reference signs and detailed description thereof will be omitted.

FIG. 9 is a cross-sectional view showing a schematic configuration around a combustor of the gas turbine according to Variant 3 of the first embodiment of the present disclosure when seen in the rotor axis direction.

FIG. 10 is a schematic diagram around the combustor of the gas turbine according to Variant 3 of the first embodiment of the present disclosure.

As shown in FIG. 9 and FIG. 10, in the gas turbine 1 according to the variant, the partition plate 20 extends to the compressor outlet port flow channel F1 in the rotor radial direction Drr, and divides the compressor outlet port flow channel F1 into a plurality of sections in the rotor circumferential direction Dcr.

In this way, by dividing the compressor outlet port flow channel F1 into independent flow channels for each of the combustors 10 by the partition plate 20, the length of the air introduction passage F2 of each of the combustors 10 can be extended by the length of the compressor outlet port flow channel F1. Accordingly, by increasing the internal pressure of the combustors 10, it is possible to further enhance an effect of suppressing reduction in supply flow rate of the compressed air A and suppressing the combustion oscillations.

<Variant 4 of First Embodiment>

Next, the gas turbine 1 according to Variant 4 of the first embodiment of the present disclosure will be described.

The same components as the above-mentioned embodiment are designated by the same reference signs and detailed description thereof will be omitted.

In the above-mentioned embodiment and variants, while the example of the combustor type, referred to as an annular cylindrical (cannular) type, in which the plurality of cylindrical combustors 10 are parallelly disposed in the rotor circumferential direction Dcr in the space in the casing 5 has been exemplarily described, there is no limitation thereto. The above-mentioned embodiment and variants may be applied to an annular combustor type.

FIG. 11 is a schematic diagram around the combustor of the gas turbine according to Variant 4 of the first embodiment of the present disclosure.

As shown in FIG. 11, in the annular combustors 10, the combustor basket 12 and the transition piece 15 are connected to each other in the rotor circumferential direction Dcr, and each of the combustors 10 has a common combustion chamber. In addition, in the annular combustors 10, the plurality of burners 13 are individually disposed in the annular common combustion chamber. The supply flow rate of the compressed air A of each of the burners 13 follows a difference between an upstream pressure and a downstream pressure of the burner, and the flow rate of the compressed air A of the burner is increased as the downstream pressure is lowered. Since the annular combustors 10 have a common combustion chamber, all the burners 13 have a common downstream pressure set value of the burners 13. However, since the work of expansion occurring in each of the burners 13 fluctuates transiently, the downstream pressures of the adjacent burners also transiently differ. For this reason, like the case of the annular cylindrical combustors 10, the partition plate 20 is disposed in the casing 5 as shown in FIG. 11, so that the air introduction passage F2, into which the compressed air A flows, is independent for each of the burners 13. Further, the partition plate 20 may be inclined with respect to the meridian plane of the annular combustors 10.

Accordingly, even in the gas turbine 1 having the annular combustors 10, it is possible to suppress deviation of the supply amount of the compressed air A and suppress the combustion oscillations.

Second Embodiment

Next, a gas turbine 1 according to a second embodiment of the present disclosure will be described.

The same components as the above-mentioned embodiment are designated by the same reference signs and detailed description thereof will be omitted.

In the related art, for the purpose of cooling the combustor, a cylindrical cooling flow channel cover may be disposed to cover the outer circumference of the combustor. Some of the air passing through the gap (cooling flow channel) between the cooling flow channel cover and the outer circumferential surface of the combustor passes through the cooling air inlet port provided in the combustor basket of the combustor, and flows into the combustor (combustor basket). In the case of such a configuration, when the combustion oscillations occur, the pressure of the combustion oscillation due to the cooling air inlet port may propagate through the air introduction passage and enlarge the combustion oscillations. Since the size of the cooling air inlet port is smaller than that of the flow channel area of the air introduction passage, the combustion oscillations with a low frequency of about 100 Hz are less likely to propagate, and it is assumed that expansion of the combustion oscillations becomes higher in a high frequency region of an order of kHz. In order to suppress the high frequency combustion oscillations, it is desirable to isolate the cooling air inlet port from the air introduction passage.

In addition, in order to increase the length of the flow channel for inertia, it is desirable to reduce the flow channel area of the air introduction passage. However, when the flow channel area is reduced, the flow velocity is increased, the pressure loss is increased, and finally, efficiency of the combustor may be decreased. In addition, the flow resistance is increased because the cooling flow channel in the related art has the internal structure (a structure provided on the outer circumferential surface of the combustor, or the like). Here, in the embodiment, the air introduction passage isolated from the conventional cooling flow channel is provided to suppress the pressure loss.

FIG. 12 is a cross-sectional view showing a schematic configuration around a combustor of the gas turbine according to the second embodiment of the present disclosure along a rotor axis.

As shown in FIG. 12, the gas turbine 1 according to the embodiment includes an annular air introduction passage F3 defined by a first cylinder section 30 that surrounds the combustors 10 and a second cylinder section 31 that surrounds the first cylinder section 30, and configured to introduce the compressed air A in the casing 5 into the combustors 10, instead of the air introduction passage F2 according to the first embodiment.

In addition, a gap between the first cylinder section 30 and the outer circumferential surface of the combustor 10 (the combustor basket 12 and the transition piece 15) functions as a cooling flow channel 32 through which some of the compressed air A circulates and configured to cool the outer circumferential surface of the combustor 10. A cooling air inlet port 33 in communication with the combustor basket 12 is provided on a downstream side of the cooling flow channel 32 (the base side Dcb of the combustor 10), and the compressed air A passing through the cooling flow channel 32 flows into the combustor 10 (the combustor basket 12) from the cooling air inlet port 33.

As described above, the gas turbine 1 according to the embodiment includes the annular air introduction passage F3 defined by the first cylinder section 30 that surrounds the combustors 10 and the second cylinder section 31 that surrounds the first cylinder section 30, and configured to introduce the compressed air A in the casing 5 into the combustors 10.

In this way, since the outer circumferential surface of the combustor 10 is covered with the first cylinder section 30, there is no internal structure that prohibits a flow of the compressed air A in an air introduction passage F3. Accordingly, the compressed air A can smoothly circulate in the air introduction passage F3.

In addition, the air introduction passage F3 is isolated from the cooling flow channel 32 by the first cylinder section 30. Accordingly, the combustion oscillations can be suppressed from being amplified by the cooling air inlet port 33 of the cooling flow channel 32.

Third Embodiment

Next, a gas turbine 1 according to a third embodiment of the present disclosure will be described.

The same components as the above-mentioned embodiment are designated by the same reference signs and detailed description thereof will be omitted.

FIG. 13 is a cross-sectional view showing a schematic configuration around a combustor of the gas turbine according to the third embodiment of the present disclosure along a rotor axis.

As shown in FIG. 13, the gas turbine 1 according to the embodiment includes an air introduction passage F4 defined by the combustors 10 with a first wall section 40 having an L-shaped cross section and a second wall section 41 having a U-shaped cross section.

The first wall section 40 is configured of a first plate section 40a extending from an outer circumferential surface of the combustor 10, and a second plate section 40b extending from the first plate section 40a to the axis downstream side Dad of the rotor 4 along the outer circumferential surface of the combustor 10.

The second wall section 41 is configured of a third plate section 41a extending from the inner circumferential surfaces 5a and 5b of the casing 5 to a downstream side Dad in the axial direction of the rotor 4 along the outer circumferential surface of the combustor 10, a fourth plate section 41b extending from the third plate section 41a toward the outer circumferential surface of the combustor 10, and a fifth plate section 41c extending from the fourth plate section 41b to the axis upstream side Dau of the rotor 4 along the outer circumferential surface of the combustor 10.

In addition, the second plate section 40b of the first wall section 40 is disposed at an interval between the third plate section 41a and the fifth plate section 41c of the second wall section 41.

The compressed air A in the casing 5 flows into the air introduction passage F4 from a space between the outer circumferential surface of the combustor 10 and the fifth plate section 41c of the second wall section 41, and flows toward the upstream side Dau in the axis direction. The compressed inflow air A in the air introduction passage F4 changes a direction and flows toward the downstream side Dad in the axial direction through a flow channel between the second plate section 40b of the first wall section 40 and the fifth plate section 41c of the second wall section 41 upon collision with the first plate section 40a of the first wall section 40. In addition, the compressed air A changes a direction and flows toward the axial direction upstream side Dau through a flow channel between the second plate section 40b of the first wall section 40 and the fourth plate section 41b of the second wall section 41 upon collision with the fourth plate section 41b of the second wall section 41. In this way, the compressed air A is introduced into the combustors 10 while changing the direction in the air introduction passage F4.

As described above, the gas turbine 1 according to the embodiment includes the air introduction passage F4 configured to change a circulation direction of the compressed air A therein using the first wall section 40 and the second wall section 41. By providing such a configuration, the flow channel length of the air introduction passage F4 can be increased. Accordingly, according to the increase in internal pressure of the combustor 10, a decrease in supply flow rate of the compressed air A can be suppressed, and the combustion oscillations can be suppressed.

Fourth Embodiment

Next, a gas turbine 1 according to a fourth embodiment of the present disclosure will be described.

The same components as the above-mentioned embodiment are designated by the same reference signs and detailed description thereof will be omitted.

FIG. 14 is a cross-sectional view showing a schematic configuration around a combustor of the gas turbine according to the fourth embodiment of the present disclosure along a rotor axis.

As shown in FIG. 14, the gas turbine 1 according to the embodiment includes a plurality of guide pipes 50. The inside of the guide pipes 50 functions as an air introduction passage F5 configured to introduce the compressed air A into the combustors 10. In addition, the guide pipes 50 are provided at an interval in the circumferential direction Dcc of the combustors 10.

Each of the guide pipes 50 has a first opening section 51 connected to an inlet port of the combustor 10 (a passage between the outer shell 11 and the combustor basket 12) and a second opening section 52 that opens in the space in the casing 5 on the downstream side Dad in the axial direction of the rotor 4.

In this way, by dividing the air introduction passage F5 into the plurality of guide pipes 50, a degree of freedom of the disposition in the casing 5 can be enhanced.

In addition, the plurality of guide pipes 50 may have different lengths. Accordingly, resonance strength of the guide pipes 50 can be weakened.

Fifth Embodiment

Next, a gas turbine 1 according to a fifth embodiment of the present disclosure will be described.

The same components as the above-mentioned embodiment are designated by the same reference signs and detailed description thereof will be omitted.

In the above-mentioned embodiment, the technology of suppressing the combustion oscillations in units of the combustors 10 has been described. In the embodiment, a technology of suppressing the combustion oscillations generated in units of the burners 13 in the combustors 10 will be described. Further, in the embodiment, while an example in which a configuration of suppressing the combustion oscillations in units of the burners 13 is applied to the configuration of the first embodiment will be described, there is no limitation thereto. In the other embodiments, the configuration of the embodiment may be applied to each of variants of the first embodiment and second to fourth variants.

FIG. 15 is a cross-sectional view showing a schematic configuration of the combustor according to the fifth embodiment of the present disclosure along a combustor axis.

FIG. 16 is a cross-sectional view showing a schematic configuration of the combustor according to the fifth embodiment of the present disclosure in the combustor axis direction.

As shown in FIG. 15, one of the combustors 10 has the plurality of burners 13. The plurality of burners 13 are disposed along the combustor axis Ac. Each of the burners 13 is configured of a fuel supply part 131 configured to supply fuel F, a nozzle 132 configured to inject the fuel from a tip thereof, and a nozzle tube 133 that surrounds the nozzle 132 concentrically.

In addition, the combustors 10 according to the embodiment further include burner partition plates 60 extending along the axis direction Dc of the combustors 10 between the adjacent burners 13 and configured to divide an outlet port space 12b of the air introduction passage F2 in the combustor basket 12 in sections independent for each other for the plurality of burners 13, respectively.

As shown in FIG. 16, the burner partition plates 60 are provided between the main burners 13b arranged in the circumferential direction Dcc. In addition, the burner partition plates 60 are provided between the pilot burners 13a and the main burners 13b to surround the pilot burners 13a.

The frequency of the combustion oscillations has an order of one hundred Hz in units of the combustors 10. On the other hand, the frequency of the combustion oscillations has an order of kHz in units of the burners 13. This depends on a difference in both dimensions.

When the introduction routes of the compressed air A of the plurality of burners 13 belonging to one of the combustors 10 are independent by providing the burner partition plate 60 in this way, each of the burners 13 has its own inertia of the compressed air independently. Then, when a pressure of a certain burner 13 is increased transiently, the inertia of the compressed air of each of the burners 13 can reduce the extent to which the compressed air bypasses the other burners 13.

As described above, it is known that abnormal combustion referred to as backfire occurs when the compressed air A is reduced. Since the backfire occurs first from the burners 13 in which the flow rate of the compressed air A is small, making the compressed air A of the burners 13 independent with respect to the pressure variation of the burners 13 is effective in preventing the backfire.

<Variant 1 of Fifth Embodiment>

Next, a gas turbine 1 according to Variant 1 of a fifth embodiment of the present disclosure will be described.

The same components as the above-mentioned embodiment are designated by the same reference signs and detailed description thereof will be omitted.

FIG. 17 is a cross-sectional view showing a schematic configuration of a combustor according to the variant of the fifth embodiment of the present disclosure along a combustor axis.

As shown in FIG. 17, in the combustors 10 according to the variant, a burner communication hole 61 bringing the sections adjacent to the outlet port space 12b in communication with each other is provided in the burner partition plate 60.

The burner communication hole 61 is formed by attaching a flow resistance material such as a perforated metal to the burner partition plate 60.

When a pressure of a certain burner 13 is higher than that of the other burners 13, a flow according to the bypass of the compressed air occurs in the burner communication hole 61. Since the bypass flow is applied as an attenuation force when the flow resistance is added to the burner communication hole 61, the pressure fluctuations can be suppressed. When the pressure of all the burners 13 belonging to one of the combustors 10 vibrates with the same phase and the same amplitude, a differential pressure does not occur on the burner partition plate 60. Accordingly, the attenuation force is not also obtained. Here, since vibrations occur in units of the combustors 10, it will be dealt with by the configurations of the first to fourth embodiments. The variant is effective against the combustion oscillations between the plurality of burners 13 in one of the combustors 10, and reduces the combustion oscillations.

Sixth Embodiment

Next, a gas turbine 1 according to a sixth embodiment of the present disclosure will be described.

The same components as the above-mentioned embodiment are designated by the same reference signs and detailed description thereof will be omitted.

FIG. 18 is a cross-sectional view showing a schematic configuration of a combustor according to the sixth embodiment of the present disclosure along a combustor axis.

As shown in FIG. 18, the combustor 10 according to the embodiment has a fuel supply pipeline 134 in communication with each of the plurality of burners 13 and configured to distribute and supply the fuel F into the burners 13.

In the above-mentioned first to fifth embodiments, a technology of suppressing the combustion oscillations by keeping the supply flow rate of the compressed air A to the combustors 10 constant has been described. In the embodiment, a configuration in which a similar technology is applied to the fuel supply to suppress the combustion oscillations will be described.

Since a combustor pressure p_B is the downstream pressure for the fuel supply pipe, the fuel supply flow rate is decreased when the combustor pressure p_B is increased. In the embodiment, the combustion oscillations are suppressed by making its degree greater than the supply of the compressed air A that is reduced. For this reason, the equivalent length of the flow channel for the inertia of the fuel supply pipe is smaller than the equivalent length of the flow channel of the air introduction passage F2. Specifically, as shown in FIG. 18, by providing the fuel supply pipeline 134 common to the plurality of burners 13 and distributing the fuel F into the burners 13 from the fuel supply pipeline 134, the flow channel area per outer diameter is increased.

In this way, by providing the fuel supply pipeline 134 common to the plurality of burners 13, the fuel supply to the burners 13 is reduced without delay when the pressure of the combustors 10 is increased, and the combustion oscillation can be suppressed because the pressure rise is eliminated.

Seventh Embodiment

Next, a gas turbine 1 according to a seventh embodiment of the present disclosure will be described.

The same components as the above-mentioned embodiment are designated by the same reference signs and detailed description thereof will be omitted.

FIG. 19 is a cross-sectional view showing a schematic configuration of a combustor according to the seventh embodiment of the present disclosure along a combustor axis.

FIG. 20 is a cross-sectional view showing a schematic configuration of the combustor according to the seventh embodiment of the present disclosure in the combustor axis direction.

As shown in FIG. 19 and FIG. 20, the combustors 10 according to the embodiment have fuel circulation pipelines 135 in communication with the plurality of burners 13, respectively, and in which the burners 13 circulate the fuel F to each other.

In the above-mentioned first to fifth embodiments, the technology of suppressing the combustion oscillations by keeping the supply flow rate of the compressed air A to the combustors 10 constant has been described. In the embodiment, a configuration of suppressing combustion oscillation by applying the same technology to the fuel supply will be described.

Since the combustor pressure p_B is the downstream pressure for the fuel supply pipe, the fuel supply flow rate decreases when the combustor pressure p_B increases. In the embodiment, the combustion oscillations are suppressed by making its degree greater than the supply of the compressed air A that is reduced. For this reason, the equivalent length of the flow channel for the inertia of the fuel supply pipe when seen from the combustor pressure p_B is smaller than the equivalent length of the flow channel of the air introduction passage F2. Specifically, as shown in FIG. 19 and FIG. 20, by connecting the fuel supply pipes of the plurality of burners 13 to the fuel circulation pipelines 135 and distributing the fuel F to the burners 13 with the low combustor pressure p_B from the burners 13 with the high the combustor pressure p_B without delay, combustion of the burners 13 with the high the combustor pressure p_B is decreased and combustion of the burners 13 with the low combustor pressure p_B is increased to uniformize combustion of the plurality of burners 13.

In this way, since the pressure of the combustors 10 is deviated and eliminated without delay by providing the fuel circulation pipelines 135 common to the plurality of burners 13, the combustion oscillations can be suppressed.

As described above, while some embodiments according to the present invention have been described, all these embodiments are provided as examples and are not intended to limit the scope of the invention. These embodiments can be implemented in various other forms, and various omissions, substitutions and modifications may be made without departing from the spirit of the present invention. These embodiments and their variants are included in the scope or spirit of the present invention as well as the scope of the present invention described in the claims and their equivalents. For example, a configuration that the plurality of combustors are disposed at equal intervals in the circumferential direction is explained in the above-described embodiments. However, the configuration of combustors is not limited thereto. In another embodiment, some of the plurality of combustors may be disposed at unequal interval in the circumferential direction. Even with such a configuration, the same effects explained in the above-described embodiments can be obtained.

<Supplementary Statement>

The gas turbines described in the above-mentioned embodiments are grasped, for example, as follows.

(1) According to a first aspect of the present disclosure, a gas turbine (1) includes a rotor (4) that is rotatable about an axis, a casing (5) configured to cover the rotor (4) in a circumferential direction and having an annular space therein, a compressor (2) configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing (5), a plurality of combustors (10) disposed at equal intervals in the casing (5) in a circumferential direction of the rotor (4) and configured to combust the compressed air and fuel taken in from the casing (5) to generate a combustion gas, a turbine (3) driven by the combustion gas, a partition plate (20) configured to divide a space in the casing (5) in the circumferential direction of the rotor (4), and an air introduction passage (F2) defined by inner circumferential surfaces (5a, 5b) of the casing (5) and configured to introduce the compressed air in the casing (5) into the combustors (10).

By providing such a configuration, when an internal pressure of a certain combustor of the plurality of combustors is increased, the gas turbine can suppress the compressed air in the casing from bypassing to the other combustors from the combustor. Accordingly, it is possible to suppress an increase and decrease in supply flow rate of the compressed air to each of the combustors due to fluctuation of the combustor pressure. In addition, since it is possible to suppress a decrease of the compressed air supplied to the combustor in which the internal pressure is increased, it is possible to suppress the combustion oscillations in the combustor and occurrence of backfire.

(2) According to a second aspect of the present disclosure, in the gas turbine (1) according to the first aspect, the partition plate (20) is provided in the middle of each of the plurality of combustors (10).

Accordingly, since the gas turbine has the air introduction passage independent for each combustor, even when an internal pressure of each combustor differs, it is possible to suppress occurrence of a difference in flow rate of the compressed air supplied to each combustor. Accordingly, it is possible to suppress the combustion oscillations more reliably.

(3) According to a third aspect of the present disclosure, in the gas turbine (1) according to the first or second aspect, a baffle plate (21) is provided in an air introduction passage (F2) and configured to bypass the compressed air.

By providing such a configuration, it is possible to lengthen the flow channel length of the air introduction passage (lengthen the equivalent length for inertia). Accordingly, by increasing the internal pressure of the combustor, it is possible to suppress a decrease in supply flow rate of the compressed air and suppress the combustion oscillations.

(4) According to a fourth aspect of the present disclosure, in the gas turbine (1) according to the second or third aspect, a communication hole (22) in communication with the air introduction passages (F2) adjacent to each other in the circumferential direction is provided in the partition plate (20).

By providing such a configuration, since the flow of the bypass of the compressed air is applied as an attenuation force, the pressure fluctuation can be suppressed.

(5) According to a fifth aspect of the present disclosure, in the gas turbine (1) according to any one of the second to fourth aspects, a compressor outlet port flow channel (F1) is connected to the casing (5) on an inner side of the rotor (4) in the radial direction, the partition plate (20) extends to the compressor outlet port flow channel (F1) in the radial direction of the rotor (4), and divides the compressor outlet port flow channel (F1) into a plurality of sections in the circumferential direction of the rotor (4).

By providing such a configuration, a length of an air circulation path of each combustor can be lengthened by a length of the compressor outlet port flow channel. Accordingly, it is possible to suppress a decrease in supply flow rate of the compressed air due to the increase in the internal pressure of the combustor and enhance an effect of suppressing the combustion oscillations.

(6) According to a sixth aspect of the present disclosure, a gas turbine (1) includes a rotor (4) that is rotatable about an axis, a casing (5) configured to cover the rotor (4) in a circumferential direction and having an annular space therein, a compressor (2) configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing (5), a plurality of combustors (10) disposed in the casing (5) at equal intervals in the circumferential direction of the rotor (4) and combust the compressed air and fuel taken in from the casing (5) to generate a combustion gas, a turbine (3) driven by the combustion gas, and an annular air introduction passage (F3) defined by a first cylinder section (30) that surrounds the combustor (10) and a second cylinder section (31) that surrounds the first cylinder section (30) and configured to introduce the compressed air in the casing (5) into the combustor (10).

In this way, since an outer circumferential surface of the combustor is covered with the first cylinder section, there is no internal structure that prohibits a flow of the compressed air in the air introduction passage. Accordingly, the compressed air can smoothly circulate in the air introduction passage.

(7) According to a seventh aspect of the present disclosure, a gas turbine (1) includes a rotor (4) that is rotatable about an axis, a casing (5) configured to cover the rotor (4) in a circumferential direction and having an annular space therein, a compressor (2) configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing (5), a plurality of combustors (10) disposed in the casing (5) at equal intervals in the circumferential direction of the rotor (4) and configured to combust the compressed air and fuel taken in from the casing (5) to generate a combustion gas, a turbine (3) driven by the combustion gas, and an air introduction passage (F4) configured to introduce the compressed air in the casing (5) into the combustor. The air introduction passage (F4) is defined by a first wall section (40) having an L-shaped cross section configured of a first plate section (40a) extending from an outer circumferential surface of the combustor (10) and a second plate section (40b) extending from the first plate section (40a) to a downstream side of the rotor (4) in an axial direction, and a second wall section (41) having a U-shaped cross section configured of a third plate section (41a) extending from inner circumferential surfaces (5a, 5b) of the casing (5) toward the downstream side of the rotor (4) in the axial direction, a fourth plate section (41b) extending from the third plate section (41a) toward an outer circumferential surface of the combustor (10) and a fifth plate section (41c) extending from the fourth plate section (41b) toward an upstream side of the rotor (4) in the axial direction. The second plate section (40b) of the first wall section (40) is disposed between the third plate section (41a) and the fifth plate section (41c) of the second wall section (41).

By providing such a configuration, a flow channel length of the air introduction passage can be increased. Accordingly, it is possible to suppress a decrease in supply flow rate of the compressed air due to the increase in the internal pressure of the combustor and suppress the combustion oscillations.

(8) According to an eighth aspect of the present disclosure, a gas turbine (1) includes a rotor (4) that is rotatable about an axis, a casing (5) configured to cover the rotor (4) in a circumferential direction and having an annular space therein, a compressor (2) configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing (5), a plurality of combustors (10) disposed in the casing (5) at equal intervals in the circumferential direction of the rotor (4) and configured to combust the compressed air and fuel taken in from the casing (5) to generate a combustion gas, a turbine (3) driven by the combustion gas, an air introduction passage (F5) defined by a plurality of guide pipes (50) having a first opening section (51) connected to an inlet port of the combustor (10) and a second opening section (52) that opens on the downstream side of the rotor (4) in the axial direction in the space in the casing (5).

In this way, by dividing the air introduction passage into a plurality of guide pipes, a degree of freedom of disposition in the casing can be increased.

(9) According to a ninth aspect of the present disclosure, in the gas turbine (1) according to any one of the above-mentioned first to eighth aspects, the combustor (10) includes a combustor basket (12) into which the compressed air is supplied from the casing (5) through the air introduction passage, a plurality of burners (13) each of which extends in the axial direction of the combustor (10), is provided at an interval along an inner circumference of a combustor basket (12) and has a nozzle (132) configured to inject the fuel from a tip and a nozzle tube (133) that concentrically surrounds the nozzles (132), and burner partition plates (60) extending between the adjacent burners (13) in the axial direction of the combustor (10) and configured to divide an outlet port space (12b) of the air introduction passage in the combustor basket (12) for each of the burners (13) independently.

By providing the burner partition plates in this way, when the introduction routes of the compressed air of the plurality of burners belonging to one combustor are provided independently, each burner has the inertia of the compressed air independently. Then, when a pressure of a certain burner is increased transiently, the inertia of the compressed air of each burner can reduce the degree to which the compressed air bypasses the other burners.

(10) According to a tenth aspect of the present disclosure, in the gas turbine (1) according to the ninth aspect, a burner communication hole (61) in communication with the adjacent sections is provided in the burner partition plate (60).

When a pressure of a certain burner is higher than the other burners, a flow according to the bypass of the compressed air occurs in the burner communication hole. Since the flow of the bypass is applied as an attenuation force when flow resistance is added to the burner communication hole, the pressure fluctuation can be suppressed.

(11) According to an eleventh aspect of the present disclosure, in the gas turbine (1) according to any one of the first to eighth aspects, the combustor (10) includes a combustor basket (12) into which the compressed air is supplied from the casing (5) through the air introduction passage, a plurality of burners

(13) each of which extends in the axial direction of the combustor (10), is provided at an interval along an inner circumference of the combustor basket (12) and has a nozzle (132) configured to inject the fuel from a tip and a nozzle tube (133) that concentrically surrounds the nozzles (132), and a fuel supply pipeline (134) connected to each of the burners (13) and configured to distribute and supply the fuel into the burners (13).

In this way, by providing the fuel supply pipeline common to the plurality of burners, even when the pressure of the combustor is deviated, the combustion oscillations can be suppressed.

(12) According to a twelfth aspect of the present disclosure, in the gas turbine (1) according to any one of the first to eighth aspects, the combustor (10) includes a combustor basket (12) into which the compressed air is supplied from the casing (5) through the air introduction passage, a plurality of burners (13) each of which extends in the axial direction of the combustor (10), is provided at an interval along an inner circumference of the combustor basket (12) and has a nozzle (132) configured to inject the fuel from a tip and a nozzle tube (133) that concentrically surrounds the nozzles (132), and a fuel circulation pipeline (135) in communication with the burners (13) and into which the fuel supplied to each of the burners (13) between the burners (13) flows.

In this way, by providing the fuel circulation pipeline common to the plurality of burners, since the pressure of the combustor is eliminated without delay even when the pressure of the combustor is deviated, the combustion oscillations can be suppressed.

According to the gas turbine according to the present disclosure, the combustion oscillations can be more effectively suppressed.

While preferred embodiments of the invention have been described and illustrated above, it should be understood that these are exemplary of the invention and are not to be considered as limiting. Additions, omissions, substitutions, and other modifications can be made without departing from the spirit or scope of the present invention. Accordingly, the invention is not to be considered as being limited by the foregoing description, and is only limited by the scope of the appended claims.

EXPLANATION OF REFERENCES

• • 1 Gas turbine
  • 2 Compressor
  • 3 Turbine
  • 4 Rotor
  • 5 Casing
  • 10 Combustor
  • 11 Outer shell
  • 12 Combustor basket
  • 13 Burner
  • 15 Transition piece
  • 20 Partition plate
  • 21, 21a, 21b, 21c Baffle plate
  • 22 Communication hole
  • 30 First cylinder section
  • 31 Second cylinder section
  • 32 Cooling flow channel
  • 33 Cooling air inlet port
  • 40 First wall section
  • 41 Second wall section
  • 50 Guide pipe
  • 60 Burner partition plate
  • 61 Burner communication hole
  • 131 Fuel supply part
  • 132 Nozzle
  • 133 Nozzle tube
  • 134 Fuel supply pipeline
  • 135 Fuel circulation pipeline
  • F1 Compressor outlet port flow channel
• F2, F3, F4, F5 Air introduction passage

Claims

1. A gas turbine comprising:

a rotor that is rotatable about an axis;

a casing configured to cover the rotor in a circumferential direction and having an annular space therein;

a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing;

a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas;

a turbine driven by the combustion gas; and

an air introduction passage defined by a partition plate configured to divide the space in the casing in the circumferential direction of the rotor and an inner circumferential surface of the casing and configured to introduce the compressed air in the casing into the combustor.

2. The gas turbine according to claim 1, wherein the partition plate is provided in the middle of each of the plurality of combustors.

3. The gas turbine according to claim 1, wherein a baffle plate configured to bypass the compressed air is provided in the air introduction passage.

4. The gas turbine according to claim 2, wherein a communication hole allowing communication between air introduction passages adjacent to each other in the circumferential direction is provided in the partition plate.

5. The gas turbine according to claim 2, wherein a compressor outlet port flow channel is connected to the casing on an inner side in the radial direction of the rotor, and

the partition plate extends to the compressor outlet port flow channel in the radial direction of the rotor, and divide the compressor outlet port flow channel into a plurality of sections in the circumferential direction of the rotor.

6. A gas turbine comprising:

a rotor that is rotatable about an axis;

a casing configured to cover the rotor in a circumferential direction and having an annular space therein;

a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing;

a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas;

a turbine driven by the combustion gas; and

an annular air introduction passage defined by a first cylinder section that surrounds the combustor and a second cylinder section that surrounds the first cylinder section and configured to introduce the compressed air in the casing into the combustor.

7. A gas turbine comprising:

a rotor that is rotatable about an axis;

a casing configured to cover the rotor in a circumferential direction and having an annular space therein;

a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing;

a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas;

a turbine driven by the combustion gas; and

an air introduction passage configured to introduce the compressed air in the casing into the combustor,

wherein the air introduction passage is defined by

a first wall section having an L-shaped cross section configured of a first plate section extending from an outer circumferential surface of the combustor and a second plate section extending from the first plate section toward a downstream side of the rotor in an axial direction, and

a second wall section having a U-shaped cross section configured of a third plate section extending from an inner circumferential surface of the casing toward the downstream side of the rotor in the axial direction, a fourth plate section extending from the third plate section toward the outer circumferential surface of the combustor and a fifth plate section extending from the fourth plate section toward the upstream side of the rotor in the axial direction, and

the second plate section of the first wall section is disposed between the third plate section and the fifth plate section of the second wall section.

8. A gas turbine comprising:

a rotor that is rotatable about an axis;

a casing configured to cover the rotor in a circumferential direction and having an annular space therein;

a compressor configured to generate a high pressure of compressed air obtained by compressing external air and send the compressed air into the casing;

a plurality of combustors disposed in the casing at equal intervals in the circumferential direction of the rotor and configured to combust the compressed air and fuel taken in from the casing to generate a combustion gas;

a turbine driven by the combustion gas; and

an air introduction passage defined by a plurality of guide pipes having a first opening section connected to an inlet port of the combustor and a second opening section that opens in the space in the casing on the downstream side in an axial direction of the rotor.

9. The gas turbine according to claim 1, wherein the combustor includes:

a combustor basket into which the compressed air is supplied from the casing through the air introduction passage;

a plurality of burners each of which extends in an axial direction of the combustor, is provided at intervals along an inner circumference of the combustor basket and has a nozzle configured to inject the fuel from a tip and a nozzle tube that concentrically surrounds the nozzle; and

a burner partition plate extending between the adjacent burners in the axial direction of the combustor and configured to divide an outlet port space of the air introduction passage in the combustor basket into independent sections for each of the burners.

10. The gas turbine according to claim 9, wherein a burner communication hole in communication with the adjacent sections is provided in the burner partition plate.

11. The gas turbine according to claim 1, wherein the combustor includes:

a combustor basket into which the compressed air is supplied from the casing through the air introduction passage;

a plurality of burners each of which extends in an axial direction of the combustor, is provided at an interval along an inner circumference of the combustor basket and has a nozzle configured to inject the fuel from a tip and a nozzle tube that concentrically surrounds the nozzle; and

a fuel supply pipeline connected to each of the burners and configured to distribute and supply the fuel to the burners.

12. The gas turbine according to claim 1, wherein the combustor includes:

a combustor basket into which the compressed air is supplied from the casing through the air introduction passage;

a plurality of burners each of which extends in an axial direction of the combustor, is provided at an interval along an inner circumference of the combustor basket and has a nozzle configured to inject the fuel from a tip and a nozzle tube that concentrically surround the nozzles; and

a fuel circulation pipeline in communication with the burners and into which the fuel supplied to each of the burners between the burners flows.