JP6866187B2 - Turbine nozzle and radial turbine equipped with it - Google Patents

Description

The present disclosure relates to a turbine nozzle and a radial turbine including the turbine nozzle.

Turbines are used to extract power from compressible working fluids such as air. The types of turbines are mainly axial flow turbines and radial turbines. Radial turbines are generally more efficient in a single stage than axial-flow turbines. Therefore, radial turbines are suitable, for example, for small to medium-sized power generation facilities.

One of the important parts of a radial turbine is the turbine nozzle. The turbine nozzle is a component for guiding the working fluid to the turbine wheel, and plays a role of expanding the working fluid and converting the pressure into a velocity. As described in Patent Document 1, in a radial turbine, a plurality of turbine vanes constituting the turbine nozzle are arranged in an annular shape around the turbine wheel. The flow path of the working fluid is formed by the space between the turbine vanes adjacent to each other in the circumferential direction of the turbine wheel. The flow path cross-sectional area is usually gradually reduced from upstream to downstream (ie, towards the turbine wheel) to expand the working fluid.

As it passes through the turbine nozzle, the working fluid expands and accelerates in response to its pressure. The impulse applied to the blade when the working fluid collides with the blade of the turbine wheel and the reaction that the working fluid expands and applies to the blade when the working fluid passes through the flow path between the blades of the turbine wheel causes the turbine wheel to It rotates (so-called impulse reaction turbine). As a result, the generator connected to the turbine wheel rotates to generate electric power.

Patent Document 2 discloses a tapered nozzle for increasing the speed of the working fluid for the purpose of increasing the output of the impulse turbine.

International Publication No. 2005/0856515 Japanese Unexamined Patent Publication No. 2010-190109 U.S. Pat. No. 5,299,909

NACA TECHNICAL NOTE No.1651 SUPESONIC NOZZLE DESIGN

One way to increase the efficiency of a radial turbine is to increase the expansion ratio of the fluid in the radial turbine. However, a radial turbine using a tapered nozzle cannot expand the working fluid at a pressure ratio (expansion ratio) exceeding the critical pressure ratio. The "critical pressure ratio" means the pressure ratio when the flow velocity of the working fluid reaches the speed of sound.

It is an object of the present disclosure to provide a technique for expanding a working fluid at a high pressure ratio exceeding the critical pressure ratio.

That is, the present disclosure is
A turbine nozzle used in radial turbines
A ring-shaped hub with a central axis and
A plurality of nozzle vanes arranged at equal intervals on the hub along the circumferential direction of the hub.
A flow path formed between the ventral surface of the nozzle vanes adjacent to each other in the circumferential direction and the back surface of the nozzle vanes, and
With
When the direction from the outer peripheral side of the hub to the inner peripheral side of the hub is defined as the flow direction of the working fluid in the flow path,
The flow path includes a throat having the smallest flow path cross-sectional area in the flow direction.
The flow path cross-sectional area is expanded on the downstream side in the flow direction from the throat.
The height of the nozzle vane on the downstream side in the flow direction from the throat is larger than the height of the nozzle vane in the throat, and gradually increases from the upstream side to the downstream side in the flow direction. Provide a turbine nozzle.

According to the technique of the present disclosure, the working fluid can be expanded at a high pressure ratio exceeding the critical pressure ratio.

FIG. 1 is a partial cross-sectional view of a radial turbine according to an embodiment of the present disclosure. FIG. 2 is a partial plan view of the radial turbine shown in FIG. FIG. 3 is a partially enlarged plan view of the turbine nozzle. FIG. 4 is an enlarged plan view of the nozzle vane. FIG. 5 is an enlarged plan view of the trailing edges of the two nozzle vanes. FIG. 6A is a cross-sectional view of the turbine nozzle along the center line of the flow path. FIG. 6B is a cross-sectional view of the turbine nozzle according to the modified example along the center line of the flow path. FIG. 6C is a cross-sectional view of the turbine nozzle according to another modification along the center line of the flow path. FIG. 7 is a graph showing the calculation result of the equation (3) when the working fluid is standard air (κ = 1.4). FIG. 8A is a graph showing the angle between the plane including the central axis and the airfoil centerline in the nozzle vane where the Mach number M of the flow velocity at the outlet of the turbine nozzle is M = 1.4. FIG. 8B is a graph showing the angle formed by the plane including the central axis and the airfoil centerline in another nozzle vane in which the Mach number M of the flow velocity at the outlet of the turbine nozzle is M = 1.4. FIG. 9 is a graph showing an example of the distribution regarding the thickness of the nozzle vanes. FIG. 10 is a graph showing the height distribution of nozzle vanes. FIG. 11 is a configuration diagram of a power generation system using a radial turbine.

(Knowledge on which this disclosure was based)
Assuming that the working fluid is an ideal fluid, the flow velocity of the working fluid at the outlet of the nozzle is expressed by the following equation (1).

Cs: Discharge flow velocity Cp: Constant pressure specific heat T ₀₁ : Static temperature at throat P _exit : Static pressure at nozzle outlet P ₀₀ : Static pressure at nozzle inlet κ: Specific heat ratio

_{The discharge flow velocity Cs is determined according to the pressure ratio P exit} / P ₀₀ , with the upper limit being the speed of sound determined according to the physical properties and the state quantity of the working fluid. The pressure ratio at which the discharge flow velocity Cs reaches the speed of sound is called the "critical pressure ratio". A general nozzle, such as a tapered nozzle, cannot expand the working fluid at a pressure ratio greater than or equal to the critical pressure ratio. That is, expansion in which the flow velocity of the working fluid exceeds the speed of sound cannot be achieved.

Next, the value M defined by the following equation (2) is called a Mach number. The Mach number is a value obtained by dividing the flow velocity by the speed of sound.

M: Mach number V: Flow velocity of working fluid a: Sound velocity κ: Specific heat ratio R: Gas constant of _{working fluid T 00} : Static temperature of working fluid

In the case of a tapered nozzle, the flow velocity is maximized at the portion where the cross-sectional area of the flow path is the minimum. When the maximum flow velocity reaches M = 1, the expansion ratio in the tapered nozzle reaches the critical pressure ratio, and the working fluid cannot expand any more. There is a relationship of the following equation (3) between the flow path cross-sectional area and the Mach number M.

A: Flow path cross-sectional area at an arbitrary position of the nozzle A *: Minimum flow path cross-sectional area of the nozzle M: Mach number κ: Specific heat ratio

FIG. 7 shows the calculation result of the equation (3) when the working fluid is standard air (κ = 1.4). As can be seen from equation (3) and FIG. 7, if the Mach number M of the flow is less than 1 at any position of the nozzle, the nozzle has the smallest flow path cross-sectional area (ie, M = 1). It is necessary to have a cross-sectional area larger than the cross-sectional area of the time. The cross-sectional area of the flow path decreases as the flow velocity increases, and the flow velocity reaches the speed of sound at the position where the cross-sectional area of the flow path is minimized. When the flow velocity exceeds the speed of sound, the cross-sectional area of the flow path expands. That is, in order to increase the flow velocity beyond the speed of sound, it is necessary to increase the cross-sectional area of the flow path.

As can be understood from these facts, in order to change the flow velocity from subsonic flow to supersonic flow, a part having a tapered shape, a part having a minimum flow path cross-sectional area (throat), and a divergent shape And a nozzle with. Nozzles with such a structure are called "Laval nozzles" and are used in propulsion engines that make heavy use of supersonic flow, such as rocket engines or aircraft engines.

In Patent Document 1, a tapered nozzle for increasing the speed of the working fluid to be guided to the turbine wheel of the impulse turbine is used for the purpose of increasing the output of the impulse turbine. The impulse turbine is configured so that the nozzle expands the working fluid almost completely and the turbine wheel is rotated by the impulse applied to the blade when the working fluid collides with the blade of the turbine wheel. The structure in which the tapered nozzle disclosed in Patent Document 1 is arranged in the tangential direction of the turbine wheel is often adopted for a turbine that operates under the conditions of a small flow rate and a high pressure ratio. However, according to this structure, the nozzle portion becomes long, so that the overall size of the turbine becomes excessive. Further, the nozzle disclosed in Patent Document 1 has a minimum flow path cross-sectional area at the tip of the nozzle. Therefore, in the nozzle disclosed in Patent Document 1, the Mach number M does not exceed 1, and acceleration in which the Mach number exceeds 1 cannot be achieved.

On the other hand, Patent Document 3 discloses a supersonic distributor for an axial flow turbine. In the supersonic distributor of Patent Document 3, the outer shape of the vane has a straight portion on the upstream side, a convex portion forming a throat, and a curved portion on the downstream side. Patent Document 3 describes that a supersonic flow having a Mach number in the range of Mach 1.2 to 2.5 can be generated. The supersonic distributor disclosed in Patent Document 3 is similar to a Laval nozzle. However, the two-dimensional shape of the flow path formed between the adjacent vanes is inevitably asymmetric with respect to the center line of the flow path due to the structural restrictions of the distributor.

On the other hand, as shown in FIG. 1 of Non-Patent Document 1, the ideal Laval nozzle is axisymmetric. According to the axisymmetric structure, the shock wave generated after passing through the throat can be reflected and canceled by the opposing wall surfaces to prevent a steep pressure change (FIGS. 8 and 9 of Non-Patent Document 1). As a result, a supersonic flow can be efficiently generated.

When the flow path does not have a symmetrical structure as in the distributor of Patent Document 3, in addition to not being able to sufficiently obtain the effect of canceling the shock wave, the flow field such as the enlargement of the boundary layer and the separation of the boundary layer. Disturbance is likely to occur. As a result, it is often possible to achieve only expansion up to a high transonic speed region with M = 1.1 to 1.2. In other words, if expansion to a higher supersonic region is required, additional ingenuity is required.

The turbine nozzle according to the first aspect of the present disclosure is
A turbine nozzle used in radial turbines
A ring-shaped hub with a central axis and
A plurality of nozzle vanes arranged at equal intervals on the hub along the circumferential direction of the hub.
A flow path formed between the ventral surface of the nozzle vanes adjacent to each other in the circumferential direction and the back surface of the nozzle vanes, and
With
When the direction from the outer peripheral side of the hub to the inner peripheral side of the hub is defined as the flow direction of the working fluid in the flow path,
The flow path includes a throat having the smallest flow path cross-sectional area in the flow direction.
The flow path cross-sectional area is expanded on the downstream side in the flow direction from the throat.
The height of the nozzle vane on the downstream side in the flow direction from the throat is larger than the height of the nozzle vane in the throat, and gradually increases from the upstream side to the downstream side in the flow direction. Is.

According to the turbine nozzle of the first aspect, the effect obtained in the Laval nozzle, for example, the effect of canceling the shock wave is improved. As a result, expansion at a higher pressure ratio can be achieved. Even after the Mach number M of the flow velocity of the working fluid reaches 1 at the throat, the working fluid can continue to accelerate, that is, continue to expand. Compared to a turbine nozzle that uses a simple tapered nozzle, a faster working fluid can be introduced into the turbine wheel, so the impulse component that rotates the turbine wheel increases, and a radial turbine that can exert a large output in a single stage. Can be built.

In the second aspect of the present disclosure, for example, the upper surface of the hub is perpendicular to the central axis and the upper surface of the nozzle vane is the center on the downstream side in the flow direction of the turbine nozzle according to the first aspect. It is tilted with respect to the plane perpendicular to the axis. According to the second aspect, processing for manufacturing a turbine nozzle is easy.

In the third aspect of the present disclosure, for example, the upper surface of the nozzle vane is perpendicular to the central axis and the upper surface of the hub is the center on the downstream side in the flow direction of the turbine nozzle according to the first aspect. It is tilted with respect to the plane perpendicular to the axis. According to the third aspect, since the upper surface of the nozzle vane is perpendicular to the plane perpendicular to the central axis of the hub, it is easy to adjust the dimension of the clearance between the nozzle vane and the shroud wall of the radial turbine. That is, it is not essential to change the shape of the shroud wall, and it is possible to suppress an increase in the manufacturing cost of the turbine nozzle.

In the fourth aspect of the present disclosure, for example, the upper surface of the nozzle vane is inclined with respect to the plane perpendicular to the central axis on the downstream side in the flow direction of the turbine nozzle according to the first aspect. Moreover, the upper surface of the hub is inclined with respect to the plane perpendicular to the central axis. According to the fourth aspect, the inclination angle of the upper surface of the nozzle vane and the inclination angle of the upper surface of the hub can be reduced.

In the fifth aspect of the present disclosure, for example, the airfoil centerline of each of the plurality of nozzle vanes of the turbine nozzle according to the first aspect includes a first portion and a second portion, and the first portion is the same. The portion from the upstream end of the airfoil centerline to the first point, the first point is the point at which the airfoil centerline begins to bend in the direction toward the central axis, and the second point. The portion is a portion from the first point to the downstream end of the airfoil centerline.

According to the fifth aspect, when the flow velocity of the working fluid reaches the supersonic speed, the direction of the shock wave generated at the trailing edge portion of the nozzle vane can be deflected to the downstream side in the flow direction. A high expansion ratio can be achieved by shifting the pressure recovery position due to the shock wave to the downstream side and expanding the region of the expansion wave generated prior to the shock wave (that is, the expansion region where the flow velocity continues to increase). Further, the inflow angle of the working fluid from the turbine nozzle to the turbine wheel can be appropriately maintained.

In the sixth aspect of the present disclosure, for example, when the angle formed by the plane including the central axis of the turbine nozzle according to the fifth aspect and the airfoil centerline is defined as an angle β, the angle is defined in the first portion. The average rate of change of β is a positive value, and the second portion includes a second point in which the average rate of change of the angle β changes from a positive value to a negative value. In the section from the point to the downstream end, the average rate of change of the angle β is a negative value. According to the sixth aspect, the distribution of the discharge speed in the width direction of the nozzle vane becomes uniform. As a result, fluctuations in angular velocity (fluctuations in torque) per rotation of the turbine wheel are suppressed, so that a generator connected to the radial turbine can generate high-quality AC power.

In the seventh aspect of the present disclosure, for example, when the angle formed by the plane including the central axis of the turbine nozzle according to the fifth or sixth aspect and the airfoil centerline is defined as an angle β, the angle β is defined as an angle β. It changes linearly in a section of a predetermined length including the downstream end of the second portion.

The radial turbine according to the eighth aspect of the present disclosure is
With any one of the turbine nozzles of the first to seventh aspects,
A turbine wheel arranged inside the turbine nozzle and
It is equipped with.

According to the eighth aspect, it is possible to provide a radial turbine capable of exhibiting a large output in a single stage.

Hereinafter, embodiments of the present disclosure will be described with reference to the drawings. The present disclosure is not limited to the following embodiments.

As shown in FIG. 1, the radial turbine 100 according to the present embodiment includes a turbine wheel 10, a shaft 12, a turbine nozzle 14, and a casing 20. The turbine wheel 10 and the turbine nozzle 14 are arranged in the casing 20. The turbine wheel 10 is arranged inside the turbine nozzle 14. The shaft 12 is fixed to the turbine wheel 10. The turbine wheel 10 includes a plurality of blades 11 and a hub 13. A plurality of moving blades 11 are provided on the surface of the hub 13 at equal angular intervals. The casing 20 includes a scroll chamber 15 and a shroud wall 16. The scroll chamber 15 is an annular space formed around the turbine nozzle 14. An intake port (not shown) provided in the casing 20 opens toward the scroll chamber 15. The working fluid is guided from the scroll chamber 15 through the turbine nozzle 14 to the turbine wheel 10. The shroud wall 16 covers the rotor blade 11 and the turbine nozzle 14 from one side in a direction parallel to the rotation axis O common to the turbine wheel 10 and the shaft 12. The rotation axis O coincides with the central axis of the turbine nozzle 14. Therefore, in the present specification, the central axis of the turbine nozzle 14 is also referred to as "central axis O".

As shown in FIG. 2, the turbine nozzle 14 is composed of a hub 22 and a plurality of nozzle vanes 24. The hub 22 is a ring-shaped and plate-shaped component. The hub 22 has a circular inner peripheral edge and an outer peripheral edge, respectively, in a plan view. The plurality of nozzle vanes 24 are arranged on the hub 22 at equal angular intervals along the circumferential direction of the hub 22.

The radial turbine 100 of the present embodiment is a so-called impulse reaction turbine. Generally, in a turbine nozzle using a nozzle vane, it is difficult to achieve expansion with a large pressure ratio because the length of each flow path is relatively short. However, according to the impulse reaction turbine, the working fluid can be first expanded at the turbine nozzle and further expanded at the turbine wheel. Since the expansion of the working fluid is shared by both the turbine nozzle and the turbine wheel, the flow velocity of the working fluid is unlikely to become excessive in each element. In this case, the impulse reaction turbine is more likely to achieve higher efficiency than the impulse turbine because friction loss and flow disturbance, which are dominated by flow velocity, can be suppressed.

As shown in FIG. 3, each nozzle vane 24 has a ventral surface 24p, a back surface 24q and an upper surface 24r. The ventral surface 24p is a surface of the hub 22 on the side close to the central axis O. The back surface 24q is a surface on the side far from the central axis O of the hub 22. In other words, the ventral surface 24p is the surface on the side closer to the turbine wheel 10, and the back surface 24q is the surface on the side far from the turbine wheel 10. The upper surface 24r is a surface facing the shroud wall 16 (see FIG. 1). The nozzle vane 24 has a columnar shape as a whole. A flow path 27 for a working fluid is formed between the ventral surface 24p of the nozzle vanes 24 adjacent to each other in the circumferential direction of the hub 22 and the back surface 24q of the nozzle vanes 24.

In the present embodiment, the flow path 27 has a contracted portion 27a, a throat 27b, and a divergent portion 27c. When the direction from the outer peripheral side of the hub 22 to the inner peripheral side of the hub 22 is defined as the flow direction of the working fluid in the flow path 27, the contracted portion 27a, the throat 27b, and the divergent portion 27c are in this order from the upstream side in the flow direction. Lined up in. The contracted portion 27a is a portion located on the upstream side in the flow direction with respect to the throat 27b, and has a channel cross-sectional area that is gradually reduced. The throat 27b is a portion having the smallest flow path cross-sectional area. The throat 27b may have a constant length in the flow direction. That is, the flow path 27 may have a section having the minimum flow path cross-sectional area. The divergent portion 27c is a portion located downstream of the throat 27b in the flow direction and has a gradually expanding channel cross-sectional area. That is, the turbine nozzle 14 of the present embodiment has a structure similar to the Laval nozzle.

As shown in FIGS. 3 and 4, in the plan view of the turbine nozzle 14, the position on the ventral surface 24p of the nozzle vane 24 corresponding to the throat 27b is defined as the specific position P1. Further, the position on the airfoil center line L of the nozzle vane 24, which is a% advanced from the upstream end Q1 of the airfoil center line L toward the downstream end Q2, is defined as the position Pa. Define. The position on the airfoil center line L, which is b% advanced from the upstream end Q1 of the airfoil center line L toward the downstream end Q2, is defined as the position Pb (a <b). Define. At this time, the intersection K of the vertical line drawn from the specific position P1 to the airfoil center line L and the airfoil center line L exists between the position Pa and the position Pb. In one example, a = 20 and b = 25 are set.

When the throat 27b is present at the position as described above, it is possible to avoid a steep reduction in the cross-sectional area of the flow path of the contracted portion 27a. As a result, excessive acceleration of the working fluid at the contracted portion 27a can be avoided. In particular, when a working fluid having a high viscosity is used, the flow path cross-sectional area of the contracted portion 27a matches the design intention, and it is possible to avoid the occurrence of flow chokes at the contracted portion 27a. Further, since the length of the divergent portion 27c that guides the flow of supersonic speed is sufficiently secured, sufficient expansion can be achieved.

According to the turbine nozzle 14 of the present embodiment, expansion at a pressure ratio exceeding the critical pressure ratio can be achieved even when expansion exceeding the critical pressure ratio is required and / or when the speed of sound in the working fluid is low. .. As a result, a large output can be obtained by a single radial turbine 100. If the temperature of the working fluid at the inlet of the turbine is low or the molecular weight of the working fluid is large, the speed of sound in the working fluid is also low.

In the present specification, the "airfoil centerline L" can be specified by the following method. First, a plan view of the nozzle vane 24 is prepared, and the chord direction is determined. The chord direction is determined so that the maximum chord length can be secured. Next, a plurality of dividing lines are drawn perpendicular to the chord direction so that the nozzle vane 24 is divided into the plurality of portions along the chord direction. By connecting the midpoints of each dividing line, the airfoil center line L is obtained. The finer the dividing line is drawn, the more accurate the airfoil centerline L can be obtained. The thickness of the nozzle vane 24 is specified by the length of a line segment that passes through an arbitrary point on the airfoil center line L and connects the ventral surface 24p and the back surface 24q at the shortest distance.

As shown in FIGS. 4 and 5 (a), the nozzle vane 24 has a main body portion 241 and a trailing edge portion 242. The trailing edge portion 242 is a portion that includes the downstream end Q2 of the airfoil center line L and is curved toward the center axis O of the hub 22. The main body portion 241 is a portion including the upstream end Q1 of the airfoil center line L and located closer to the upstream end Q1 of the airfoil center line L than the trailing edge portion 242. As shown in FIG. 5A, the airfoil centerline L of the nozzle vane 24 includes a first portion L1 and a second portion L2. The first portion L1 is a portion from the upstream end Q1 of the airfoil center line L to the first point B. The first point B is a point at which the airfoil center line L begins to bend in the direction toward the center axis O. The second portion L2 is a portion from the first point B to the downstream end Q2 of the airfoil center line L. In the present embodiment, the point B is a boundary point between the trailing edge portion 242 and the main body portion 241 on the airfoil center line L. According to such a structure, when the flow velocity of the working fluid reaches the supersonic speed, the direction of the shock wave generated in the trailing edge portion 242 can be deflected to the downstream side in the flow direction. A high expansion ratio can be achieved by shifting the pressure recovery position due to the shock wave to the downstream side and expanding the region of the expansion wave generated prior to the shock wave (that is, the expansion region where the flow velocity continues to increase). Further, the inflow angle of the working fluid from the turbine nozzle 14 to the turbine wheel 10 can be appropriately maintained.

When it is intended to expand the working fluid at a large pressure ratio, the region of the expansion wave tends to be terminated by the shock wave (pressure wave) generated at the trailing edge portion of the nozzle vane in the Laval nozzle or the nozzle similar to the Laval nozzle. On the other hand, according to the present embodiment, the region of the expansion wave can be expanded to the downstream side of the trailing edge portion 242 of the nozzle vane 24. Therefore, it is possible to expand the working fluid with a larger pressure ratio. As a result, the working fluid having a higher flow velocity flows into the turbine wheel 10 from the turbine nozzle 14. Since the impulse for driving the turbine wheel 10 is increased, the output of the radial turbine 100 is improved. Further, since the flow velocity distribution in each flow path 27 is smoothed, the fluctuation of the angular velocity (torque fluctuation) per rotation of the turbine wheel 10 is suppressed, and the waveform of the generated AC power approaches the sinusoidal waveform. That is, high quality power can be obtained. Since the working fluid is guided from the turbine nozzle 14 toward the turbine wheel 10 at an appropriate angle, the heat insulation efficiency of the radial turbine 100 is also improved.

As shown in FIG. 5A, it is a position on the airfoil center line L and advances by x% of the total length of the airfoil center line L from the upstream end Q1 of the airfoil center line L toward the downstream end Q2. The position is defined as the position Px. Similarly, the position on the airfoil center line L, which is y% ahead of the total length of the airfoil center line L from the upstream end Q1 of the airfoil center line L toward the downstream end Q2, is the position Py (b <. It is defined as x <y). The boundary point B between the trailing edge portion 242 and the main body portion 241 on the airfoil center line L exists, for example, between the position Px and the position Py. In one example, x = 85 and y = 90 are set. According to such a structure, it is possible to form an expanded expansion region without inhibiting the expansion in the divergent portion 27c. As a result, the output of the radial turbine 100 is improved.

FIG. 8A is a graph showing a change in the angle β formed by the plane including the central axis O and the airfoil center line L in the nozzle vane 24 having the trailing edge portion 242 of the shape shown in FIG. 5 (a). The horizontal axis represents the ratio of the distance from the upstream end Q1 of the airfoil center line L to the total length of the airfoil center line L. The vertical axis represents the angle β at each position on the airfoil center line L. As can be seen from FIG. 8A, the average rate of change of the angle β is not constant. According to such a structure, the pressure fluctuation due to the shock wave (compression wave) generated in the trailing edge portion 242 is linearly directed toward the downstream side at an angle determined according to the angle of the trailing edge portion 242. It occurs between the nozzle vanes 24 and the nozzle vanes 24. In the expanded expansion region, the distribution of the discharge velocity in the width direction of the nozzle vane 24 becomes uniform. As a result, fluctuations in angular velocity (fluctuations in torque) per rotation of the turbine wheel 10 are suppressed, so that a generator connected to the radial turbine 100 can generate high-quality AC power.

The above-mentioned effect can be enhanced by increasing the degree of bending (bending) at the boundary point B. The trailing edge portion 242 of the nozzle vane 24 shown in FIG. 5 (b) is significantly curved at the boundary point B as compared with the trailing edge portion 242 of the nozzle vane 24 shown in FIG. 5 (a). FIG. 5 (c) shows the trailing edge portion of the nozzle vane shown in FIG. 5 (a) and the trailing edge portion of the nozzle vane shown in FIG. 5 (b) superimposed for comparison.

FIG. 8B is a graph showing the change in the angle β formed by the plane including the central axis O and the airfoil center line L in the nozzle vane 24 having the trailing edge portion 242 of the shape shown in FIG. 5 (b). As can be understood from FIG. 8B, the airfoil center line L of the nozzle vane 24 having the shape of FIG. 5 (b) is greatly curved at the second point C slightly closer to the downstream end Q2 than the boundary point B. .. In the first portion L1 of the airfoil center line L, the average rate of change of the angle β is a positive value. The second portion L2 of the airfoil centerline L includes a second point C in which the average rate of change of the angle β changes from a positive value to a negative value. At the second point C, the angle β changes from monotonically increasing to monotonically decreasing. In other words, at the second point C, the average rate of change of the angle β changes from a positive value to a negative value. In the section from the second point C to the downstream end Q2, the average rate of change of the angle β is a negative value. According to such a structure, the above-mentioned effect can be further enhanced. In the present embodiment, the boundary point B is different from the second point C. However, the boundary point B may coincide with the second point C.

As shown in FIG. 8B, the angle β changes linearly in a section of a predetermined length including the downstream end Q2. In the section from the second point C to the downstream end Q2, the average rate of change of the angle β is substantially constant (the slope is constant). When the trailing edge portion 242 has such a structure, the expansion region can be expanded as described above. Since the restriction is applied so that the discharge angle of the working fluid from the turbine nozzle 14 is not excessively deflected, the inflow angle of the working fluid into the turbine wheel 10 can be maintained at an appropriate value as designed. As a result, the efficiency of the radial turbine 100 is further improved.

In the present embodiment, the thickness of the nozzle vane 24 gradually decreases from a position slightly downstream of the intersection K described above. Specifically, as shown in FIG. 4, it is a position on the airfoil center line L, and is c% of the total length of the airfoil center line L from the upstream end Q1 to the downstream end Q2 of the airfoil center line L. The advanced position is defined as the position Pc (b <c <x). In the example of FIG. 4, the thickness of the nozzle vane 24 starts to decrease from an arbitrary position included in the section from the position Pb to the position Pc toward the downstream end Q2 of the airfoil center line L. In one example, c = 30 is set. According to such a change in thickness, the throat 27b can be formed at an appropriate position.

FIG. 9 is a graph showing an example of the distribution regarding the thickness of the nozzle vane 24 of FIG. The horizontal axis represents the ratio of the distance from the upstream end Q1 of the airfoil center line L to the total length of the airfoil center line L. The vertical axis represents the ratio of the distance from the airfoil center line L to the surface of the nozzle vane 24 in the thickness direction of the airfoil vane 24 with respect to the total length of the airfoil center line L. In FIG. 9, the solid line shows the ratio regarding the distance (first thickness) from the airfoil center line L to the back surface 24q in the thickness direction of the nozzle vane 24. In FIG. 9, the broken line shows the ratio regarding the distance (second thickness) from the airfoil center line L to the ventral surface 24p in the thickness direction of the nozzle vane 24. The thickness of the nozzle vane 24 at an arbitrary position on the airfoil center line L is represented by the sum of the first thickness and the second thickness. In the example of FIG. 9, the thickness of the nozzle vane 24 is about 20% of the total length of the airfoil center line L from the upstream end Q1 of the airfoil center line L toward the downstream end Q2, that is, when a = 20. It becomes the maximum at the position Pa or the position before and after it. The thickness of the nozzle vane 24 at a position slightly downstream from the intersection K existing between the position Pa when a = 20 and the position Pb when b = 25 (for example, the position Pc when c = 30). It clearly shows a decreasing trend. The thickness of the nozzle vane 24 decreases monotonically and gradually from the position Pc to the downstream end Q2. According to such a change in thickness, the throat 27b can be formed at an appropriate position.

In the present specification, the dimension of the nozzle vane 24 with respect to the direction parallel to the central axis O of the hub 22, and the dimension from the upper surface 22p of the hub 22 to the upper surface 24r of the nozzle vane 24 is defined as the height of the nozzle vane 24. The height of the nozzle vane 24 on the downstream side in the flow direction from the throat 27b is larger than the height of the nozzle vane 24 on the upstream side in the flow direction from the throat 27b. According to such a structure, the effect obtained in the Laval nozzle, for example, the effect of canceling the shock wave is improved. As a result, expansion at a higher pressure ratio can be achieved. Even after the Mach number M of the flow velocity of the working fluid reaches 1 at the throat 27b, the working fluid can continue to accelerate, that is, continue to expand. Compared to a turbine nozzle that uses a simple tapered nozzle, a faster working fluid can be introduced into the turbine wheel 10, so the impulse component that rotates the turbine wheel 10 increases, and a radial that can exert a large output in a single stage. The turbine 100 can be constructed.

Specifically, as shown in FIGS. 6A to 6C, the height H of the nozzle vane 24 gradually increases from the upstream side to the downstream side in the flow direction on the downstream side in the flow direction from the throat 27b. The height H of the nozzle vane 24 on the downstream side in the flow direction from the throat 27b is larger than the height h of the nozzle vane 24 on the upstream side in the flow direction from the throat 27b. According to such a structure, the change in the cross-sectional area of the flow path can be brought close to that of the Laval nozzle. As a result, the working fluid can be expanded more smoothly.

6A to 6C are cross-sectional views of the nozzle vane 24 along the center line of the flow path 27 shown in FIG. The ventral surface 24p of the nozzle vane 24 appears in FIGS. 6A to 6C.

In the example shown in FIG. 6A, the upper surface 22p of the hub 22 is perpendicular to the central axis O of the hub 22 and the upper surface 24r of the nozzle vane 24 is perpendicular to the central axis O of the hub 22 on the downstream side in the flow direction from the throat 27b. It is tilted with respect to the plane. In the example shown in FIG. 6A, the thickness of the hub 22 in the direction parallel to the central axis O is constant. Since the thickness of the hub 22 is constant, processing for producing the shape shown in FIG. 6A is easy.

In the example shown in FIG. 6B, the upper surface 24r of the nozzle vane 24 is perpendicular to the central axis O of the hub 22 and the upper surface 22p of the hub 22 is perpendicular to the central axis O of the hub 22 on the downstream side in the flow direction from the throat 27b. It is tilted with respect to the plane. In this example, the thickness of the hub 22 varies along the nozzle vanes 24. The thickness of the hub 22 is reduced on the downstream side of the throat 27b in the flow direction. According to the example shown in FIG. 6B, since the upper surface 24r of the nozzle vane 24 is perpendicular to the plane perpendicular to the central axis O, the dimension of the clearance between the nozzle vane 24 and the shroud wall 16 (see FIG. 1) is adjusted. Is easy. That is, it is not essential to change the shape of the shroud wall 16, and it is possible to suppress an increase in the manufacturing cost of the turbine nozzle 14.

In the example shown in FIG. 6C, the upper surface 24r of the nozzle vane 24 is inclined with respect to the plane perpendicular to the central axis O of the hub 22 on the downstream side in the flow direction from the throat 27b. Further, the upper surface 22p of the hub 22 is inclined with respect to a plane perpendicular to the central axis O of the hub 22. This example is a combination of the example of FIG. 6A and the example of FIG. 6B. According to the example shown in FIG. 6C, the inclination angle of the upper surface 24r of the nozzle vane 24 and the inclination angle of the upper surface 22p of the hub 22 can be reduced.

FIG. 10 is a graph showing the height distribution of nozzle vanes. The horizontal axis represents the ratio of the distance from the upstream end Q1 of the airfoil center line L to the total length of the airfoil center line L. The vertical axis represents the ratio of the height of each position to the maximum height. The height of the nozzle vane 24 at each position represents the height on the airfoil center line L. In this embodiment, the nozzle vane 24 has the maximum height at the downstream end Q2. The height of the nozzle vane 24 is constant from the upstream end Q1 (0% position) to the position Pb. As described with reference to FIGS. 3 and 4, the position Pb is a position slightly downstream of the intersection K. The height of the nozzle vane 24 increases substantially linearly from the position Pb to the downstream end Q2 (100% position). According to such a structure, a sudden pressure change is suppressed on the downstream side of the throat 27b, and the working fluid can be expanded more smoothly.

In the present embodiment, the starting point of the divergent portion 27c is on the downstream side of the throat 27b. In this embodiment, the throat 27b has a certain length. That is, the turbine nozzle 14 of the present embodiment has a section having the minimum flow path cross-sectional area. In one example, the throat 27b has a length of about 5% of the total length of the airfoil centerline L. The starting point of the divergent portion 27c is set at the position of the downstream end of the throat 27b. Since a boundary layer is formed on the surface of the nozzle vane 24, the flow of the working fluid is narrowest at a position downstream of the head position of the throat 27b. In consideration of this, the starting point of the Suehiro portion 27c is determined. This change in the cross-sectional area of the flow path is given by the shape of the airfoil center line L of the nozzle vane 24, the thickness of the ventral surface 24p side of the nozzle vane 24, the thickness of the back surface 24q side of the nozzle vane 24, and the height of the nozzle vane. .. This is due to the fact that the thickness of the nozzle vane 24 decreases from the position Pb slightly downstream of the intersection K and that the height of the nozzle vane 24 increases from the position Pb slightly downstream of the intersection K. ing.

Next, an embodiment of a power generation system using the radial turbine 100 will be described.

As shown in FIG. 11, the power generation system 200 according to the present embodiment includes a Rankine cycle circuit 110, a heat source 112, and a duct 114. The Rankine cycle circuit 110 includes a radial turbine 100, a condenser 102 (condenser), a pump 104, and an evaporator 106 (steam generator). The radial turbine 100, the condenser 102, the pump 104, and the evaporator 106 are connected in this order by a plurality of pipes. A generator 108 is connected to the rotating shaft of the radial turbine 100. When the working fluid is expanded by the radial turbine 100, the generator 108 is driven to generate electric power. The Rankine cycle circuit 110 may include other known equipment such as a reheater.

Evaporator 106, and a working fluid circulating in the heat transfer flow body 116 and the Rankine cycle circuit 110 that is generated by the heat source 112 is heat exchanger is configured to vaporize the working fluid. In this embodiment, the evaporator 106 is arranged in the duct 114. The duct 114 is connected to the heat source 112. The heat transfer fluid 116 generated by the heat source 112 flows through the duct 114. The heat transfer fluid 116 may be a gas or a liquid. When the heat transfer fluid 116 is a gas, the evaporator 106 may be composed of a gas-liquid heat exchanger such as a fin tube heat exchanger. When the heat transfer fluid 116 is a liquid, the evaporator 106 may be composed of a liquid-liquid heat exchanger such as a plate heat exchanger or a double tube heat exchanger.

The type of heat source 112 is not particularly limited. Examples of the heat source 112 include boilers, factory equipment, engines, waste incinerators, solar ponds, fuel cells and the like.

The type of working fluid of the Rankine cycle circuit 110 is also not particularly limited. The working fluid may be an organic substance such as a hydrocarbon or a halogenated hydrocarbon, or an inorganic substance such as water, ammonia or carbon dioxide. Examples of the hydrocarbon include propane and the like. Examples of the halogenated hydrocarbon include R410a, R22, R32, and R245fa.

The techniques disclosed herein are useful for radial turbines. Radial turbines are useful, for example, in power generation systems.

10 Turbine wheel 11 Drive blade 12 Shaft 13 Turbine wheel hub 14 Turbine nozzle 16 Shroud wall 22 Turbine nozzle hub 22p Hub top surface 24 Nozzle vane 24p Abdominal surface 24q Back surface 24r Top surface 27 Flow path 27a Shrinkage 27b Throat 27c Radial turbine 102 Condenser (condenser)
104 Pump 106 Evaporator (steam generator)
108 Generator 110 Rankine cycle circuit 112 Heat source 114 Duct 200 Power generation system O Rotating axis L Airfoil centerline

Claims (6)

A turbine nozzle used in radial turbines
A ring-shaped hub with a central axis and
A plurality of nozzle vanes arranged at equal intervals on the hub along the circumferential direction of the hub.
A flow path formed between the ventral surface of the nozzle vanes adjacent to each other in the circumferential direction and the back surface of the nozzle vanes, and
With
When the direction from the outer peripheral side of the hub to the inner peripheral side of the hub is defined as the flow direction of the working fluid in the flow path,
The flow path includes a throat having the smallest flow path cross-sectional area in the flow direction.
The flow path cross-sectional area is expanded on the downstream side in the flow direction from the throat.
The height of the nozzle vane on the downstream side in the flow direction from the throat is larger than the height of the nozzle vane in the throat, and gradually increases from the upstream side to the downstream side in the flow direction .
The throat has a constant length in the flow direction .
Turbine nozzle. In the plan view of the turbine nozzle, the position on the ventral surface of the nozzle vane corresponding to the downstream end of the throat is defined as the specific position P1, which is the position on the airfoil centerline of the airfoil vane and is the airfoil centerline. Position Pa is defined as a position advanced by a% of the total length of the airfoil centerline from the upstream end to the downstream end, and the total length of the airfoil centerline from the upstream end to the downstream end of the airfoil centerline. When the position advanced by b% is defined as the position Pb, the intersection K between the vertical line drawn from the specific position P1 to the airfoil center line and the airfoil center line exists between the position Pa and the airfoil center line. The turbine nozzle according to claim 1, wherein a = 20 and b = 25. The first aspect of the present invention, wherein the upper surface of the hub is perpendicular to the central axis and the upper surface of the nozzle vane is inclined with respect to a plane perpendicular to the central axis on the downstream side of the throat in the flow direction. Turbine nozzle. The first aspect of the present invention, wherein the upper surface of the nozzle vane is perpendicular to the central axis and the upper surface of the hub is inclined with respect to a plane perpendicular to the central axis on the downstream side of the throat in the flow direction. Turbine nozzle. On the downstream side of the throat in the flow direction, the upper surface of the nozzle vane is inclined with respect to the plane perpendicular to the central axis, and the upper surface of the hub is inclined with respect to the plane perpendicular to the central axis. The turbine nozzle according to claim 1, which is inclined. The turbine nozzle according to any one of claims 1 to 5 and
A turbine wheel arranged inside the turbine nozzle and
With a radial turbine.

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