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1 6. Gas Turbines 6. Gas Turbines 1 / 140

2 Fundamentals of Gas Turbines 2 Compressor 19 Combustor 40 Turbine Gas Turbines 2 / 140

3 7FA Gas Turbine for Power Generation VIGV Air Extraction Ports Transition Piece Diffuser Starter & Gear Box Air Inlet Compressor Combustor Turbine Exhaust Cold Section Hot Section 6. Gas Turbines 3 / 140

4 Gas Turbine In a gas turbine, the working fluid for transforming thermal energy into rotating mechanical energy is the hot combustion gas, hence the term gas turbine. The first power generation gas turbine was introduced by ABB in It was a standby unit with a thermal efficiency of 17%. The gas turbine technology has many applications. The original jet engine technology was first made into a heavy duty application for mechanical drive purposes. Pipeline pumping stations, gas compressor plants, and various modes of transportation have successfully used gas turbines. While the mechanical drive applications continue to have widespread use, the technology has advanced into larger gas turbine designs that are coupled to electric generators for power generation applications. Gas turbine generators are self-contained packaged power plants. Air compression, fuel delivery, combustion, expansion of combustion gas through a turbine, and electricity generation are all accomplished in a compact combination of equipment usually provided by a single supplier under a single contract. The advantages of the heavy-duty gas turbines are their long life, high availability, and slightly higher overall efficiencies. The noise level from the heavy-duty gas turbines is considerably less than gas turbines for aviation. 6. Gas Turbines 4 / 140

5 Idealized Brayton Cycle [1/3] Air 1 2 Fuel Combustor 3 Exhaust gas 4 Power Compressor Turbine p 2 Q in 3 T (h) 3 W out Q in W in W out 2 4 W in 1 Q out 4 1 Q out s 6. Gas Turbines 5 / 140

6 Idealized Brayton Cycle [2/3] Compression Process (1 2) The entering air is compressed to higher pressure. No heat is added. However, compression raises the air temperature so that the discharged air has higher temperature and pressure. The mechanical energy transmitted from the turbine is used to compress the air. Combustion Process (2 3) Compressed air enters the combustor, where fuel is injected and combustion occurs. Combustion occurs at constant pressure. However, pressure decreases slightly in the practical process. Although high local temperatures are reached within the primary combustion zone (approaching stoichiometric conditions), the combustion system is designed to provide mixing, burning, dilution, cooling. Combustion mixture leaves with mixed average temperature. The chemical energy contained in the fuel is converted into thermal energy. 6. Gas Turbines 6 / 140

7 Idealized Brayton Cycle [3/3] Expansion Process (3 4) The thermal energy contained in the hot gases is converted into mechanical work in the turbine. This conversion actually takes place in two steps: Nozzle: the hot gases are expanded and accelerated, and a portion of the pressure energy is converted into kinetic energy. Bucket: a portion of the kinetic energy is transferred to the rotating buckets and converted into mechanical work. Some of the work produced by the turbine is used to drive the compressor, and the remainder is used to drive load equipment, such as generator, ship propeller, and pump, etc. Typically, more than 50% of the work produced by the turbine section is used to power the compressor. Exhaust Process (4 1) This is a constant-pressure cooling process. This cooling is done by the atmosphere, which provides fresh, cool air as well. The actual cycle is an open rather than closed. 6. Gas Turbines 7 / 140

8 Simple Cycle Heat and Work 754 MJ/s (100%) 272 MJ/s (36.1%) 205 MW (27.2%) MW = 482 MW (63.9%) 277 MW (Net Output) (36.7%) 6. Gas Turbines 8 / 140

9 Terminology Terminology Combined cycle Simple cycle Heavy duty gas turbines Aeroderivative gas turbines Mechanical drive gas turbines Meaning Brayton cycle (topping cycle) + Rankine cycle (bottoming cycle) Gas turbine + Steam turbine HRSG links gas turbine and steam turbine The term simple cycle is used to distinguish this configuration from the complex cycles, which utilizes additional components, such as heat exchanger for regeneration, intercooler, reheating system, or steam boilers. In general. it means gas turbines for power generation because they differ from aeronautical designs in that the frames, bearings, and blading are of heavier construction. The aero-engines transformed into land based gas turbines successfully. P&W JT8/FT8, GE J79/LM1500, GE CF6/LM2500, CF6/LM5000, CF6/LM6000 The LM2500 has been the most commercially successful one. Sometimes, it includes heavy duty gas turbines, aeroderivative gas turbines, gas (oil) pumping gas turbines, and gas turbines for marine applications. Generally, this means the industrial gas turbines that are used solely for mechanical drive or used in collaboration with a recovery steam generator differ from power generating sets in that they are often smaller and feature a "twin" shaft design as opposed to a single shaft. The power range varies from 1 MW up to 50 MW. 6. Gas Turbines 9 / 140

10 Combined Cycle Power Plants [1/6] In simple cycle mode, the gas turbine is operated alone, without the benefit of recovering any of energy in the hot exhaust gases. The exhaust gases are sent directly to the atmosphere. In combined cycle mode, the gas turbine exhaust gases are sent into HRSG. The HRSG generates steam that is normally used to power a steam turbine. Fuel Combustor HP drum LP drum Exhaust gas Steam turbine G G Turbine HRSG Condenser Inlet air Compressor HP superheater HP evaporator HP economizer LP superheater LP evaporator LP economizer LP boiler feed pump HP boiler feed pump Deaerator Condensate pump 6. Gas Turbines 10 / 140

11 Combined Cycle Power Plants [2/6] Cycle Diagram for a 3 Pressure Reheat Cycle (F-Class Gas Turbine) Fuel G Gas turbine Heat recovery steam generator Air Hot reheat steam Main steam Cold reheat steam IP steam LP steam G Steam turbine Condenser Condensate pump Steam Water Fuel Air 6. Gas Turbines 11 / 140

12 Combined Cycle Power Plants [3/6] T-s Diagram for a Typical CCPPs Combined cycle power plants have a higher thermal efficiency because of the application of two complementary thermodynamic cycles T Combustion (heat In) Topping cycle Bottoming cycle Stack (heat out) Condenser (heat out) s 6. Gas Turbines 12 / 140

13 Combined Cycle Power Plants [4/6] Generals [1/2] Combined cycle power plant means a gas turbine operated with the Brayton cycle, is combined with a heat recover steam generator and steam turbine operated with the Rankine cycle, in one plant. When two cycles are combined, the efficiency increases higher than that of one cycle alone. Thermal cycles with the same or with different working fluid can be combined. In general, a combination of cycles with different working fluid has good characteristics because their advantages can complement one another. Normally, when two cycles are combined, the cycle operating at the higher temperature level is called as topping cycle. The waste heat is used for second process that is operated at the lower temperature level, and is called as bottoming cycle. The combination used today for commercial power generation is that of a gas topping cycle with a water/steam bottoming cycle. In this case heat can be introduced at higher temperature and exhausted at very low temperature. Temperature of the air used as a working fluid of gas turbines can be increased very high under lower pressure. Water/steam used as a working fluid can contain very high level of energy at lower temperature because it has very high specific heat. Normally the topping and bottoming cycles are coupled in a heat exchanger. 6. Gas Turbines 13 / 140

14 Combined Cycle Power Plants [5/6] Generals [2/2] Air is used as a working fluid in gas turbines having high turbine inlet temperatures because it is easy to get and has good properties for topping cycle. Steam/water is an ideal material for bottoming cycle because it is inexpensive, easy to get, non-hazardous, and suitable for medium and low temperature ranges. The initial breakthrough of gas-steam cycle onto the commercial power plant market was possible due to the development of the gas turbine. In the late 1970s, EGT reached sufficiently high level that can be used for high efficiency combined cycles. The breakthrough was made easier because gas turbines have been used for power generation as a simple cycle and steam turbines have been used widely. For this reason, the combined cycle, which has high efficiency, low installation cost, fast delivery time, had been developed easily. 6. Gas Turbines 14 / 140

15 Improved Condenser Heat Balance of CCPPs Combined Cycle Power Plants [6/6] Three pressure reheat cycle Fuel energy 100% Loss in HRSG 0.3% Loss 0.5% GT 37.6% ST 21.7% 8.6% Stack Loss 0.3% 31.0% 6. Gas Turbines 15 / 140

16 Simple Cycle Simple cycle gas turbines for electricity generation are typically used for standby or peaking capacity and are generally operated for a limited number of hours per year. Peaking operation is often defined as fewer than 2,000 hours of operation per year. In mechanical drive applications, and for some industrial power generation, simple cycle gas turbines are base-load and operate more than 5,000 hours of operation per year. Some plants are initially installed as simple cycle plants with provisions for future conversion to combined cycle. Gas turbines typically have their own cooling, lubricating, and other service systems needed for simple cycle operation. This can eliminate the need to tie service systems into the combined cycle addition and will allow continued operation of the gas turbine during the conversion process and, with proper provisions, during periods when the combined cycle equipment is out of service. If future simple cycle is desired, a bypass stack may be included with the connection of the HRSG. A typical method for providing this connection is to procure a divert damper box at the outlet of the gas turbine. [ with Bypass Stack ] [ without Bypass Stack ] 6. Gas Turbines 16 / 140

17 Terminology MS7001F, GE Heavy duty gas turbines LM6000, GE Aeroderivative gas turbines Newly designed for power generation High aspect ratio (long, thin) turbine blades with tip shrouds to dampen vibration and improve blade tip sealing characteristics Single-shaft Electrical output of up to 340 MW Standardized Manufactured on the base of sales forecasts rather than orders received Series of frame sizes - shorter installation time - low costs Derived from jet engines (lightweight components, compact design, and high efficiency) and frequently incorporating a separate power turbine Low aspect ratio turbine blades with no shroud Two- or three-shaft turbine with a variable speed compressor (This is an advantage for part-load efficiency because airflow is reduced at low speeds) Higher part load efficiency because of variable speed Two-shaft turbines are usually used for compressor or pump drives The size is limited to 100 MW due to the maximum size of aircraft 6. Gas Turbines 17 / 140

18 Terminology MS7001E, GE Hot-end drive MS7001F, GE Cold-end drive In the hot-end drive configuration, the output shaft extends out the rear of the turbine. The designer is faced with many constraints, such as output shaft length, high EGT, exhaust duct turbulence, pressure drop, and maintenance accessibility. Insufficient attention to any of these details, in the design process, often results in power loss, vibration, shaft or coupling failures, and increased down-time for maintenance. This configuration is difficult to service as the assembly must be fitted through the exhaust duct. In the cold-end drive configuration, the output shaft extends out the front of the compressor. The single disadvantage is that the compressor inlet must be configured to accommodate output shaft. The inlet duct must be turbulent free and provide uniform, vortex free, flow over the all operating range. Inlet turbulence may induce surge in the compressor resulting in complete destruction of the unit. 6. Gas Turbines 18 / 140

19 Fundamentals of Gas Turbines Compressor Combustor Turbine 6. Gas Turbines 19 / 140

20 Configuration of a Compressor 7EA, GE 6. Gas Turbines 20 / 140

21 Structure A stage consists of a rotor/stator combination. 6. Gas Turbines 21 / 140

22 Introduction to Compressor If pressure rise is small and mass flow is large, the device is a called a fan, whereas if the pressure rise is high, the device is called a compressor. Sometimes a middle-range pressure rise device is termed a blower. The compressors in most gas turbine applications, especially units over 5 MW, use axial flow compressors. The axial compressor is the most complicated component to design in an aerodynamic point of view. An axial flow compressor is one which the flow enters the compressor in an axial direction (parallel with the axis of rotation), and exits from the gas turbine also in an axial direction. The axial flow compressor consumes around 50% of the power produced by the turbine section of the gas turbine. The increase in gas turbine efficiency is dependent on four basic parameters: pressure ratio, TIT, compressor efficiency, and turbine efficiency. In an axial flow compressor, air passes from one stage to the next, each stage raising the pressure slightly. However, by producing low-pressure increases on the order of 1.05:1 to 1.3:1, very high compressor efficiency can be obtained. The use of multiple stages permits overall pressure increases up to 40:1. The industrial gas turbine has been conservative in the pressure ratio and TIT. This is because the industrial gas turbines given up high performance for both rugged operation and long life. However, this has all changed in the last 10 years. The performance of the industrial gas turbines improved dramatically to overcome the increased energy cost. In addition, the performance gap between aerospace engines and industrial ones reduced dramatically. 6. Gas Turbines 22 / 140

23 Pressure ratio Compressor Growth of Pressure Ratio Compressor pressure ratio of a gas turbine engine is an extremely important design parameter. Pressure ratio = total outlet pressure total inlet pressure In general, the higher the pressure ratio, the greater thermal efficiency. The growth of both the pressure ratio and TIT parallel each other, as both growths are necessary to achieving the increase in thermal efficiency in gas turbines The expression compression ratio is not used because this is a ratio of air density rather than air pressure by definition. Aircraft Industrial Currently, some engines have compressor pressure ratio of 23:1 (40:1 for aircraft gas turbines) Year 6. Gas Turbines 23 / 140

24 Structure VIGV Variable inlet guide vanes, those are variable pitch blades, are frequently used at the compressor inlet to ensure that air enters the first stage rotor at the desired flow angle. Additionally, its position affects the quantity of compressor inlet air flow. 6. Gas Turbines 24 / 140

25 Structure VIGV When designing a compressor with the reaction blades, the first stage must be preceded by variable inlet guide vanes to provide pre-swirl and the correct velocity entrance angle to the first stage rotor. This means that VIGVs serve to direct the axially approaching flow correctly into the first row of rotor blades because those are very sensitive to any small change in incidence in flow angle or non-uniformity in velocity. Additionally, its position affects the quantity of compressor inlet air flow. Therefore, VIGVs are one of the useful tools to control stall occurred in compressors. IGVs also serve the purpose of preventing the injection of foreign objects into the engine. Similar vanes, often known as the EGVs (Exit Guide Vanes) are placed at the compressor exit to remove the rotational moment imparted to the air during compression. Engine center line VIGV 1 st Stage compressor blade 6. Gas Turbines 25 / 140

26 Rotor and Stator [1/3] 6. Gas Turbines 26 / 140

27 Rotor and Stator [2/3] Stator buildup Rotor buildup Compressor rotor stacking 6. Gas Turbines 27 / 140

28 Rotor and Stator [3/3] The axial compressor is a multi-stage unit as the amount of pressure rise by each stage is small; a stage consists of a row of rotor blades followed by a row of stator vanes. The entering air is accelerated in the rotor, that is, kinetic energy is transferred to the air, and then it is diffused in the stator to convert the kinetic energy into a pressure rise. The job of the rotors is to increase pressure using mechanical energy transmitted from turbine. Stage The stator vanes are placed to the rear of the rotor blades to receive the air at high velocity and act as a diffuser, converting kinetic energy to pressure energy. The stator also have a secondary function of directing airflow to the next stage of compression as the desired angle. From the front to rear of the compressor, i.e. from the low to high pressure end, there is a gradual reduction of annulus area between to rotor hub and the stator casing. This is necessary to maintain a near constant air axial velocity as the density increases through the length of the compressor. The convergence of the air annulus is achieved by the tapering of the casing or rotor. A combination of both is also possible, with the arrangement being influenced by manufacturing problems and other mechanical design factors. 6. Gas Turbines 28 / 140

29 Direction of Rotation Cascade The compressor is composed of several rows of airfoil cascades. That is, several blades placed in a row is called as cascade. The high pressure zone air of the first stage blade being pumped into the low pressure zone of its stator. The high pressure zone of the first stage stator vane then pumps into the low pressure zone of the second stage rotor blade. This cascade progress continues to the last stage of compression. It might appear that the rotor blade high and low pressure zone might cancel each other out as they blend together; but the overall effect of the divergent shape of the flow path results in a net decrease in velocity and an increase in static pressure. H L H L H L H L H L H L H L H L H L H L H L H L IGVs First Stage Rotor Blades High Pressure Stator Low Pressure Second Stage Rotor blades Stator blades 6. Gas Turbines 29 / 140

30 Compressor Blade Nomenclature [1/3] C 2 s c a 1 Point of max. camber i c 1 1 = blade inlet angle 2 = blade exit angle 1 = air inlet angle 2 = air exit angle c 1 = air inlet velocity c 2 = air exit velocity C = chord s = pitch (or space) = blade camber angle = 1 2 = deflection = 1 2 = stagger or setting angle i = incidence angle = 1 1 = deviation angle = 2 2 = solidity (= C/s) AR = aspect ratio (= h/c) 6. Gas Turbines 30 / 140

31 Compressor Blade Nomenclature [2/3] Nomenclature Description Camber line A line drawn halfway between the two surfaces, pressure and suction Camber The distance between the camber line and the chord line Camber angle, The turning angle of the camber line (= 1 2 ) Blade shape Aspect ratio, AR Pitch (blade spacing) Solidity, The blade shape is described by specifying the ratio of the chord to the camber at some particular length on the chord line, measured from the leading edge Aspect ratio is the ratio of the blade length (height) to the chord length The term hub-to-tip ratio is used frequently instead of aspect ratio It is important when 3-D flow characteristics are discussed The pitch of a cascade is the distance between blades, usually measured between the camber lines at the leading edges or trailing edges of the blades The ratio of the chord length to the pitch is the solidity of the cascade ( = C/s) It measures the relative interference effects of one blade with another : isolated airfoil test data can be applied with considerable accuracy : isolated airfoil test data can be applied with reduced accuracy : cascade data are necessary (majority of present designs belong to) 1.5 : channel theory can be employed Blade inlet angle, The angle formed by a line drawn tangent to the forward end of the camber line and the axis of 1 the compressor Blade exit angle, The angle formed by a line drawn tangent to the rear of the camber line and the axis of the 2 compressor 6. Gas Turbines 31 / 140

32 Compressor Blade Nomenclature [3/3] Nomenclature Stagger angle (setting angle ), Description The angle formed by a chord line and the axis of the compressor It is also called as setting angle of the blade High aspect ratio blades often are pretwisted so that at full speed centrifugal forces acting on the blades will untwist the blades to the designed angle. The pretwist angle at the tip for blades with AR of about 4 is between 2-4. Absolute air inlet angle, 1 Angle between absolute incoming velocity and axial direction. Absolute air exit angle, 2 Angle between absolute leaving velocity and axial direction. Relative air inlet angle, 1 Angle between relative incoming velocity and axial direction. (not shown in the figure) Relative air exit angle, 2 Angle between relative leaving velocity and axial direction. (not shown in the figure) Incidence angle, i The difference between the blade inlet angle and air inlet angle (i = 1 1 ) Angle of attack, The angle between the inlet air direction and the blade chord ( = 1 ) Deviation angle, As the air is turned by the blade, it offers resistance to turning and leaves the blade at an angle, greater than 2, so called air exit angle, 2 The deviation angle is defined as the difference between the blade or the camber angle and the average flow angle ( = 2 2 ) Deflection, The angle formed by air inlet angle and air exit angle ( = 1 2 ) 6. Gas Turbines 32 / 140

33 Rotor Stator Velocity Triangles in Axial Flow Compressors [1/4] r Flow direction CL z 6. Gas Turbines 33 / 140

34 Velocity Triangles in Axial Flow Compressors [2/4] Most axial compressors are designed on the basis of constant axial velocity throughout the stages because of the simplifications of design procedure of the subsequent stage. c : absolute velocity w : relative velocity u : tangential velocity of blade IGV Fluid velocity is an important variable governing the flow and energy transfer within a turbine. The absolute velocity ( ) is the fluid velocity relative to some stationary point and is usually parallel to the nozzle (stationary blade). When considering the flow across a rotating element like a bucket, the relative velocity ( w ) is important and is usually parallel to the rotating element. Vectorially, the relative velocity is defined as: where u of the bucket. w c u c is the tangential velocity rotor w 2 2 c 2 2 v z,2 c,2 u 2 w 1 1vz,1 stator u 1 c 3 Shaft CL 3 c 1 1 c,1 6. Gas Turbines 34 / 140

35 The air approaches the rotor blades with an absolute velocity c 1 at an inlet angle 1 to the axial direction. In combination with the tangential velocity of the rotor blades u, its relative velocity will be w 1 at an inlet angle 1. (w 1 = c 1 u) The relative velocity w 1 should align closely with the rotor blade angle at the inlet. After passing through the diverging passages formed between two adjacent rotor blades which do work on the air and increase its absolute velocity (from c 1 to c 2 ), the air will have relative velocity w 2 at an exit angle 2 which is less than 1. This turning of the air towards the axial direction is necessary to provide the increase of effective flow area. The rotor blade turns the relative velocity w 1 to w 2, thereby imparting angular momentum to the air and thus increasing the absolute tangential velocity. For a rotor, c 2 c 1 and w 2 w 1. This is because the kinetic energy is added by the shaft in the absolute frame, and rotor blade passage acts like a diffuser in relative frame. The fact of c 2 c 1 can be explained by the fact that the mechanical energy transmitted from the turbine will be transferred to the air through the rotor and the absolute velocity of the air increases. The exit relative velocity w 2 is nearly parallel to the blade at exit. The relative velocity w 2 in combination with u gives the absolute velocity c 2 at exit of the rotor at an angle 2. Velocity Triangles in Axial Flow Compressors [3/4] The absolute velocity at rotor exit should line up with the stator blades. The air then pass through the passages formed by the stator blades wherein it is further diffused to velocity c 3 at an exit angle 3. In most designs, it is equal to 1 so that it is prepared for entry to the next stage. (c 3 =c 1, and 3 = 1 ) 6. Gas Turbines 35 / 140

36 Velocity Triangles in Axial Flow Compressors [4/4] The basic principle of acceleration of the working fluid, followed by diffusion to convert acquired kinetic energy into a pressure rise, is applied in the axial compressor. Air is turned through the proper angle by the VIGVs before it impinges on the rotor blade of the first stage. Work transmitted from the turbine is added to the air by the rotor blades, thereby increasing its stagnation enthalpy, pressure, temperature, and kinetic energy. The flow is discharged at a proper angle of attack to stator blades where the static pressure is further increased by flow diffusion. IGV Rotor 1 Stator 1 c 1 w 1 u c 2 w 2 u h o, p o, T o h, p, T c 3 w 3 u Rotor 2 Shaft CL The stagnation pressure remains nearly the same through the stator (except for losses), but the static pressure is further increase while the kinetic energy decreases. The air is directed to the second stage rotor, and the process repeats itself. 6. Gas Turbines 36 / 140

37 Rotor Rotor Rotor IGV Stator Stator Stator Variation of Velocity and Pressure An axial flow compressor compresses its working fluid by first accelerating the fluid and then diffusing it to obtain a pressure increase. The fluid is accelerated by a row of rotor and diffused by a row of stator. The diffusion in the stator converts the velocity increase obtained in the rotor to a pressure increase. C L h o, p o, T o : total values h, p, T: static values Even though the pressure is rising dramatically, the velocity is held relatively constant. c: absolute velocity Compressor exiting velocity is lower than compressor entering velocity for flame stability in combustion chambers. ho 1 h c 2 2 c p T o 1 c 1 c c T 2 2 p To T 2 2c p It should be noted that total conditions for pressure, temperature, and enthalpy increase only in the rotor where energy is inputted to the system. 6. Gas Turbines 37 / 140

38 Pressure Ratio The length of the blades and the annulus area, which is the area between the shaft and shroud, decreases throughout the length of the compressor. This reduction in flow area compensates for the increase in fluid density as it is compressed, permitting a constant axial velocity. In the heavy duty gas turbines, pressure ratio per stage is reduced to provide stable operation. For example, GE s H gas turbine has a pressure ratio per stage of In the multistage compressor, the pressure ratio is obtained by multiplying the all pressure ratio per stage. (18 stages at 1.19 per stage gives a factor of ). 6. Gas Turbines 38 / 140

39 Types of Compressor Type Pressure Ratio Industrial Aviation Research Efficiency (%) Operational Range (Surge to choke) Centrifugal Large, 25% Axial Narrow, 3-10% Hub Shroud Outflow Inflow Axis of Rotation Stage A Centrifugal Compressor An Axial Flow Compressor 6. Gas Turbines 39 / 140

40 Fundamentals of Gas Turbines Compressor Combustor Turbine 6. Gas Turbines 40 / 140

41 Combustor Air Inlet Compressor Combustors Turbine Exhaust Cold Section Hot Section The function of the combustor is to add heat energy to the flowing gases, thereby expanding and accelerating the gases into the turbine section. From the viewpoint of thermodynamics, if the fuel heat is added at constant pressure, the volume of the gas is increased and, with flow area remaining the same, this causes an acceleration of gas to occur. 6. Gas Turbines 41 / 140

42 연소기일반사항 [1/2] 가스터빈을포함한모든열기관은공통적으로연료가가지고있는화학적에너지를연소를통해열에너지로변환시킨후최종적으로기계적인에너지로변환시켜동력을얻음 그런데열에너지는힘의요소가없기때문에열에너지를직접기계적인에너지로변환시킬수없음 따라서열기관은공기와같은작동유체에열에너지를공급하여작동유체의에너지수준을충분히높인후작동유체의상태변화를이용하여기계적인에너지로변환시켜동력을얻음 따라서모든열기관은작동유체의에너지수준을높이기위한연소기구비 그런데각종열기관에서연소를조절하는것은매우어려움 우선적으로대부분의열기관에사용하는화석연료는약 1800~2000 C 의연소온도에서화염이안정적으로형성되는데, 이렇게높은온도는연소기구성품소재의용융온도를훨씬초과. 따라서연소기주요구성품은내열소재로제작해야하며, 필수적으로냉각을고려해야함 가스터빈연소기는대량의공기가고속으로흘러가기때문에안정적인연소를유지시키기매우어려움 아울러이렇게높은온도에서는대표적인환경오염원인물질인 NO x 가많이발생하기때문에연소과정에서발생을억제시켜야함 왕복엔진과같은밀폐계열기관에서연소가일어나면온도상승과더불어압력도상승 그러나가스터빈과같은개방계열기관은연소가일어나면서압력상승이일어나지않도록설계해야함 만약에연소과정에서압력이상승하면화염이역류하여개방계열기관의우수성인연속운전이이루어지지않음 6. Gas Turbines 42 / 140

43 연소기일반사항 [2/2] 그런데다행스럽게도가스터빈연소기는개방되어있는공간에위치하고있으며, 개방된공간에서연소가일어나면압력상승이일어나지않음 이론적으로가스터빈연소기에서는압력이일정한상태에서연소가일어난다고가정하지만실제적으로는유체마찰과연소에의한급격한온도상승으로인해압력이약간낮아짐 연소기에서압력이낮아지면가스터빈출력과열효율이모두낮아짐. 따라서연소기는연소과정에서발생하는작동유체의압력손실이최소가되도록설계해야함 연소기는압축기및터빈과같은주변구성품과크기면에서도조화를이루어야하기때문에매우작음 따라서연소기는제한된내부공간에서최대한많은열량을발생시킬수있도록설계되어야함 압력이일정한상태에서연소가일어나면연소가스온도가상승함에따라연소가스비체적이증가하는데, 연소기출구유동단면적은항상일정하게유지되기때문에작동유체의가속이일어남 따라서연소기는큰압력에너지, 열에너지, 운동에너지를가진작동유체를터빈으로공급해주는역할수행 그리고이렇게높은수준의에너지를포함하고있는작동유체는터빈을구동시켜압축기와발전기를구동하기위한동력생산 연소기에공급하는연료량은연소기에서올려야하는작동유체의온도상승크기에의해서결정. 예를들면, 터빈블레이드소재에의해서결정되는사이클최고온도가 1500 C 이며, 압축기에서의공기압축에의하여압축공기온도가이미 500 C 가되었다면연소를통해서압축공기온도를 1000 C 상승시키면됨. 따라서연료량은압축공기온도를 1000 C 상승시킬수있는양을공급하면됨 6. Gas Turbines 43 / 140

44 연소기구성및공기흐름 Primary air: combustion ( 연료와반응하여고온가스생성 ) Secondary air: cooling ( 연소기및고온가스냉각 ) 6. Gas Turbines 44 / 140

45 디퓨저 디퓨저는압축기를빠져나온고속의압축공기속도를줄여서연소기로보내줌 일반적으로압축기를빠져나오는압축공기는마하수 0.5 정도의속도를가짐. 그런데압축공기가이렇게빠른속도로연소기에공급되면화염이안정적으로형성되기어려움 따라서압축기를빠져나온압축공기를디퓨저를통과시켜마하수 0.35 정도로감속시켜줌 연소기에서화염이안정적으로유지되기위해서는연소기입구에서공기속도가느려야함 디퓨저출구에서일반적으로나타나는마하수 0.35 의속도는매우빠른편임. 그러나디퓨저에서마하수 0.35 이하의속도로낮추기어려움 이는디퓨저에서공기속도를더낮추기위해서는디퓨저의확산각도를증가시켜야하는데, 이경우디퓨저에서유동박리가일어나기쉬우며, 유동박리가일어나면디퓨저기능을상실할뿐만아니라난류강도가증가하여압력손실이증가하는문제가발생하기때문임 따라서디퓨저에서감속시킬수있는한계는마하수 0.35 정도이며, 압축공기속도는연소기입구에서추가적으로낮추어줌 Diffuser 6. Gas Turbines 45 / 140

46 연료분사기 [1/2] 가스터빈연료는최종적으로연료분사기를통해연소영역으로분사 가스터빈에사용하는액체연료는연소시키기위해먼저기화시켜야함 연료분사기 (fuel injector) 는액체연료가쉽게기화될수있도록액체연료를무화시켜작은액적으로만들어표면적을증가시킴 액적상태의연료는가스터빈연소기에서와같이높은압력의공기속에서기화가촉진되어쉽게연소 연료분사기종류 : 압력무화식, 공기충돌식, 증발식, 예혼합식 1) 압력무화식 (pressure-atomizing) 연료분사기 연료에 500 psi 의높은압력을가하여연료를무화시킴 압력무화식연료분사기는매우간단하다는장점을가지고있지만몇가지단점도가지고있다. 첫째, 연료계통이고압에견딜수있도록견고하게제작해야하는데, 이는연료계통무게증가의원인이된다. 둘째, 연료무화가고르게일어나지않는경향을가지는데, 이는더많은공해물질과연기 (smoke) 를발생시키는불완전연소또는고르지못한연소의원인이된다. 2) 공기충돌식 (air blast) 연료분사기 얇은막형태로분사되는연료 (sheet of fuel) 에고속으로공기를충돌 (blast) 시켜균일한액적으로연료를무화시킴 연료를무화시키기위하여단지 1 차공기의일부분사용 공기충돌식연료분사기를적용하여최초로연소기에서연기가발생하는문제해결 공기충돌식연료분사기는압력무화식에비해낮은연료압력사용 6. Gas Turbines 46 / 140

47 연료분사기 [2/2] 3) 증발식연료분사기 공기충돌식과유사하며, 연료가연소영역으로분사되면서 1 차공기와혼합 그러나연료 - 공기혼합기는연소영역내부에있는튜브를따라흘러감 이과정에서연소영역의열이연료 - 공기혼합기로전달되어연소되기전에연료의일부가증발하며, 이로인해연료와공기의혼합이촉진 증발식연료분사기를이용하면복사열이줄어든상태에서연소가일어나기때문에연소기라이너를보호할수있음 그러나증발튜브는그내부를흘러가는연료에의해냉각되기때문에연료유량이적은경우에수명이짧아지는심각한문제를가지고있음 4) 예혼합 (premixing) 연료분사기 연료가연소영역에도달하기전에공기와미리혼합하여연소영역에공급 연소가일어나기전에연료와공기를미리혼합시키면연료와공기를매우균일하게혼합시킬수있으며, 이를통해공해물질발생을줄일수있음 연료와공기를미리혼합하면연료 - 공기혼합기를더낮은온도에서연소시킬수있기때문에 NO x 발생을크게줄일수있음 예혼합연료분사기의단점은미리혼합된연료 - 공기혼합기가연소영역에도달하기전에자동점화 (auto-ignition) 될수있다는점 아울러미리혼합된연료 - 공기혼합기의속도가화염전파속도보다작은경우연료 - 공기혼합기가연소영역에도달하기전에화염이역방향으로전파되어연소. 이를플래시백 (flashback) 이라함 자동점화와플래시백이일어나면연소기는심각한손상을입게됨 6. Gas Turbines 47 / 140

48 선회기 [Swirler] 1 연료와공기의혼합촉진 2 화염길이를짧게유지시켜연소실길이축소. 이를통해가스터빈무게절감 3 연료노즐하류에재순환영역을만들어화염정착을도모함으로써연소기내부에화염이안정적으로유지되게해줌 6. Gas Turbines 48 / 140

49 점화기 [Igniter] 대부분의가스터빈은자동차에사용하는점화기와유사한전기불꽃점화기 (electrical spark igniter) 사용 점화기는연료와공기가혼합되어있는연소영역에위치해야하지만연소에의해손상받지않도록연소기상류에위치해야함 점화기에의해일단연소가일어나면연소기내부에형성되는재순환영역에의해화염이유지되기때문에점화기는더이상필요하지않음 따라서점화기는점화가완료된후에집어넣을수있는형태 (retractable type) 로되어있는것도있음 6. Gas Turbines 49 / 140

50 케이싱 케이싱은연소기외부를구성하는얇은판으로서매우단순한구성품 케이싱은내부를흐르는공기에의해연소기에서발생하는열로부터보호되기때문에열적성능은전혀문제되지않음 그러나케이싱내부는높은압력이작용하는데반해외부는낮은압력이작용하기때문에연소기케이싱은일종의압력용기 따라서연소기는구조적으로안정돼야함 아울러연소기케이싱은가스터빈외관을결정하는요소이기때문에압축기및터빈케이싱과조화를이루어야함 6. Gas Turbines 50 / 140

51 라이너 [1/7] 라이너는연소기에형성되는화염을둘러싸고있는일종의튜브 따라서라이너는화염관 (flame tube) 이라고도불림 라이너에는연소기내부로 1 차공기 (primary air) 와 2 차공기 (secondary air) 를안정적으로공급하기위하여다공판형태로많은구멍이뚫려있음 6. Gas Turbines 51 / 140

52 라이너 [2/7] 연소기내부공기흐름및연소영역 Primary zone Intermediate zone Dilution zone Mixing and combustion Completion of combustion (2CO+O 2 2CO 2 ) Cooling A representative flow pattern in an annular combustor with double air swirler (from Tanica, 2000) 6. Gas Turbines 52 / 140

53 라이너 [3/7] 연료가분사되는연료노즐바로하류에 1 차영역 (primary zone) 형성 1 차영역에서는연료와공기의혼합이일어나면서연소가일어남 연료와혼합되어연소에사용되는공기를 1 차공기라함 1 차공기의일부는선회기를통해연소기중앙부로공급되며, 일부는라이너상류에있는작은구멍을통해반경방향으로공급. 1 차공기양은압축기를빠져나온전체공기의약 1/3 이하정도 2 차영역 (secondary zone, or intermediate zone) 에서는계속해서연소가진행 가스터빈연료는탄소와수소가결합되어있는탄화수소. 탄소는연소과정에서일차적으로산화되어일산화탄소를형성. 이렇게생성된일산화탄소는 2 차영역에서다시산소와결합하여이산화탄소가되면서탄소의연소완료. 따라서탄소는폭발적으로연소하는수소에비해연소가느리게진행. 그러므로연료의연소는 2 차영역에서종료되며, 1 차영역과 2 차영역을합친길이는연소기전체길이의약 75% 를차지 연소기에서생성된배기가스온도는터빈으로직접보내기에는너무높음 따라서 2 차영역하류에있는희석영역 (dilution zone, or tertiary zone) 에추가적으로압축공기를공급하여배기가스와혼합시켜배기가스온도를터빈에적합한온도로낮추어줌 희석공기는라이너하류에있는구멍을통해연소기로유입 최근에개발된가스터빈은성능을향상시키기위하여증가된 TIT 를적용함에따라더적은양의희석공기를사용. 이렇게줄어든희석공기는다시연료의연소에사용할수있기때문에가스터빈성능더욱향상되며, 추력증가. 6. Gas Turbines 53 / 140

54 라이너 [4/7] 라이너는연소기에서생성된화염을감싸고있기때문에냉각필요 라이너는냉각공기를이용하여냉각시킴. 연소기에서희석공기와냉각공기를합쳐서 2 차공기라함 라이너는장기간높은화염온도에노출되기때문에초합금 (superalloy) 으로제작 그러나초합금을이용하여제작하더라도라이너는냉각이필요하며, 이를위해연소기를빠져나온압축공기를이용하여충돌냉각과막냉각또는침출냉각을실시하며, 라이너내부에열차폐코팅을실시하기도함 공기를이용한라이너냉각에막냉각또는침출냉각방법사용 막냉각은라이너외부에있는냉각공기를라이너에있는작은구멍을통해라이너바로안쪽으로주입하는방법. 이렇게주입된냉각공기는라이너를높은열로부터보호하기위하여라이너안쪽표면에얇은냉각공기막을형성 침출냉각은라이너를다공성물질로제작하여라이너를냉각시키는방법 침출냉각을이용하면적은양의냉각공기를이용하여막냉각과비슷한냉각효과를얻을수있음 침출냉각과막냉각의큰차이점두가지는냉각결과나타나는라이너의온도분포와냉각에사용되는냉각공기양임 침출냉각은다공성물질로부터무수히많은작은구멍을통해냉각공기가균일하게주입되기때문에막냉각에비해훨씬균일한라이너온도분포를얻을수있음 막냉각의경우냉각공기는구멍이나루버를통해라이너내부로유입되기때문에구멍과구멍사이는상대적으로온도가높고, 구멍주위는온도가낮아서라이너온도분포가균일하지못함 더욱중요한것은침출냉각은막냉각에비해훨씬적은양의냉각공기를사용한다는점임 막냉각이전체공기의 20~50% 를냉각에사용하는데반해침출냉각은 10% 정도만을사용 냉각공기양을줄이면더많은공기를연소에사용할수있기때문에가스터빈성능이향상되고엔진추력도증가 6. Gas Turbines 54 / 140

55 라이너 [5/7] 막냉각 [Film Cooling] Air film flow prevent carbon from forming on the inside of the liner. Carbon deposits can cause hot spots or block cooling air passages. 6. Gas Turbines 55 / 140

56 라이너 [6/7] 막냉각및충돌냉각 충돌냉각 (impingement cooling) 은냉각공기를라이너에고속으로충돌시켜냉각시키는방법 공기를고속으로물체표면에충돌시키면대류열전달계수가증가하여물체를더욱효과적으로냉각시킬수있음 충돌냉각에사용된냉각공기는라이너에연속적으로형성되어있는슬롯을따라흐르면서냉각공기막을형성하여이차적으로막냉각이일어나게함 6. Gas Turbines 56 / 140

57 라이너 [7/7] 열차폐코팅 열차폐코팅 (TBC; Thermal Barrier Coating) 이란고온가스로부터라이너모재로전달되는열을줄여주기위해열전도도가매우낮은세라믹을이용하여라이너내부표면에얇은두께로코팅함으로써절연층을만들어주는것 이렇게라이너내부표면에절연층이형성되어있으면라이너표면을고온의배기가스로부터절연시켜라이너표면온도를낮게유지시켜줄수있음 이경우라이너온도를낮추기위한냉각공기양을줄일수있으며, 이로인해가스터빈성능이향상되고엔진추력도증가 열차폐코팅은두개의코팅층, 즉본드코팅과탑코팅으로이루어짐 본드코팅은라이너의산화저항성과부식저항성을향상시키기위해실시 Hot gases Thermal Barrier Coating Bond Coat C 고온의배기가스열이라이너로전달되는것을차단하기위해실시하는탑코팅재료는세라믹 Base Material Cooling gases 6. Gas Turbines 57 / 140

58 연소기구비조건 1) 모든운전조건에서연소효율이높아야한다. 2) 공해물질배출이최소가되어야한다. 3) 연소기에서발생하는압력손실이작아야한다. 4) 모든운전조건에서연소가안정적으로이루어져야한다. 5) 거친연소나맥동없이연소가부드럽게일어나야한다. 6) 터빈수명을보장하기위해온도변화가작아야한다. 급격한온도구배는연소기라이너비틀림과크랙의원인이되기때문에반드시피해야한다. 7) 연소기길이와직경이가스터빈외관형상에적합해야한다. 8) 제작비가저렴해야하며, 유지보수가쉬워야한다. 9) 어떤운전환경에서도카본디포짓 (carbon deposits) 이형성되지말아야한다. 카본디포짓이형성되면연소기라이너가비틀어질수있으며, 연소기내부유동패턴을달라지기때문에압력손실이증가한다. 6. Gas Turbines 58 / 140

59 연소강도 연소강도는연소기크기를결정하는요소로서요구되는열방출률, 연소실체적, 연소압력등에의해영향을받으며, 다음과같이정의 combustion intensity = heat release rate combustor vol. * pressure [kw / m 3 -atm] 열방출률은연소기를통과하는공기의질량유량, 연료 / 공기비, 연료의발열량을모두곱해서구해짐 연소강도가낮아지면앞서살펴본모든설계요구조건을만족시킬수있는연소기설계가더욱쉬워짐 이는연소기체적이증가할수록연소기에서발생하는압력손실이줄어들고, 연소효율이상승하며, 연소기출구온도분포가향상되며, 연소가더욱안정적으로일어나기때문임 액체연료를사용하는경우연소압력이증가할수록연료액적의증발이가속 아울러연소압력은연소가진행되는화학반응에큰영향을미침 따라서연소압력이증가할수록연소강도가낮아짐 항공용가스터빈연소기의연소강도는 2~5x10 4 kw/m 3 atm 정도인데반해산업용은항공용의약 1/10 정도 이는항공용가스터빈은관통유동형태의환형연소기를사용하는데반해서산업용은역류형의캔형연소기를사용하기때문에훨씬더큰연소기체적을확보할수있기때문 6. Gas Turbines 59 / 140

60 연소효율 가스터빈연소기가갖추어야할가장중요한요소가운데하나는높은연소효율로서연소효율은연소기에서발생한열량을연소기에공급한열량으로나눈값 연소효율이낮다는것은일부연료가연소되지않은상태로연소기를빠져나간다는것을의미하기때문에경제적으로큰손실이며, 연소되지않고대기중으로배출되는연료는공해물질이기때문에연소기는반드시높은연소효율을갖추어야함 연소기성능을측정하는세가지중요한요소는연소효율, 연소실에서의압력손실, 연소기출구온도분포균일도이다. 연소효율은다음과같이정의 Combustion efficiency = Fuel burnt in the combustor Total fuel input = Actually released energy Theoretically available energy 따라서연소효율은연소기에서미연소연료량을측정하는요소이기때문에연료소모율에영향을주는중요한설계요소 과거에는연소기를설계하는데있어서가장중요한목표가높은연소효율을얻는것과연기 (visible smoke) 발생을줄이는것이었는데, 이두가지모두 1970 년대초반에해결 현재대부분의가스터빈연소효율은 99% 임. 한편, 일산화탄소와미연탄화수소배출규제를만족시키기위해서탈설계조건에서연소효율은반드시 98.5% 를초과해야함 6. Gas Turbines 60 / 140

61 Pressure loss factor 압력손실 연소기에서발생하는압력손실은가스터빈효율과출력에큰영향을미치기때문에연소기설계에있어가장중요한요소가운데하나임 연소기에서는정압을기준으로대개 2~4% 압력손실이발생하는데, 가스터빈효율은연소기에서발생하는압력손실과동일한퍼센트만큼낮아짐 Fundamental pressure loss 연소기에서는유체마찰, 난류유동, 그리고연소에의한온도상승으로인해서압력손실이발생 10 Cold loss, K 1 유체마찰에의해발생하는압력손실을마찰손실또는냉손실 (cold loss) 연소에의한온도상승으로인해발생하는압력손실을기본손실 (fundamental loss) 또는온손실 (hot loss) 이라함 이가운데마찰손실이기본손실에비해훨씬큰비중차지 T o,2 PLF K 1 K2 1 To, Temperature ratio, T o,2 /T o,1 V 1 V p o,1 po,2 To,2 2 / 2 1V 1 To, Gas Turbines 61 / 140

62 연소기출구온도분포 연소기설계에영향을미치는요소가운데하나인연소기출구온도분포는원주방향과반경방향으로구분 Tip 가스터빈연소기출구에서의원주방향온도분포균일도는버켓수명에큰영향을미침. 이는연소기를빠져나오는연소가스의원주방향온도분포가균일하지못하면원주방향으로회전하는버켓의수명이열피로에의해크게단축될수있기때문임 한편, 온도분포균일도가나빠지면터빈블레이드크립손상에영향을미치는첨두온도 (peak temperature) 가 TIT 에비해커짐 따라서연소기출구평균온도는첨두온도에의해제한되며, 이로인해가스터빈성능을지배하는요소인 TIT 는연소기출구온도분포에의해영향을받음 Temperature Hub 연소기하류에있는터빈버켓은매우높은온도환경에서운전되며, 고속회전에따른원심력때문에높은응력이발생하기때문에뛰어난크립특성을지닌소재로제작해야함 그런데원심력에의해발생하는응력은버켓팁에서허브쪽으로가면서증가하여허브부분에서가장커지기때문에버켓의크립손상을최소화하기위해연소기출구온도분포는반경방향을따라제어되어야함 일반적으로연소기출구첨두온도는반경방향을따라서허브로부터버켓길이의 2/3 지점에나타나도록함 6. Gas Turbines 62 / 140

63 공해물질배출 가스터빈연소실에서연소가진행되는과정에서원치않는공해물질발생 현재법적으로배출이규제되고있는네가지공해물질 : 미연탄화수소 ( 미연소연료 ), 연기 (smoke, 탄소입자 ), 일산화탄소, NOx 이런공해물질생성에영향을미치는요소는연소기내부에형성되는압력및온도와연소시간 연소기 1 차영역의연료농후영역에서는연료가연소하면서일산화탄소와연기가생성되며, 이들은 2 차영역에서계속해서연소되어독성이없는이산화탄소가됨 미연탄화수소역시 2 차영역에서완전연소시켜발생을줄일수있음 NOx 는연소에의한고온환경에서생성되며, 고온에서의체류시간이길어질수록발생이증가 따라서 NOx 발생을줄이기위해서는가능한빨리화염을냉각시키는것이바람직하며, 연소시간을짧게유지해야함 현재는연소기설계기술의지속적인향상으로과거에비해공해물질배출이훨씬줄어들었음 6. Gas Turbines 63 / 140

64 캔형연소기 다중연소기 (multiple chamber) 이라고도불리는캔형연소기는하나의가스터빈에여러개의연소기가터빈과압축기를연결하는축주위에반경방향으로배치되며, 각각의연소실로압축기를빠져나온공기가덕트를통해유입 비록하나의가스터빈에여러개의연소기가설치되지만서로마주보는두개의연소기에만점화플러그가설치되며, 나머지연소기에는각각의연소실을서로연결해주는인터콘넥터튜브 (interconnector tube, or interconnector) 를통해화염이전파되어연소가일어남 또한인터콘넥터튜브는각각의연소실이동일한압력에서운전되도록해주기때문에가스터빈진동을줄여주는역할수행 인터콘넥터튜브는크로스파이어튜브 (crossfire tube) 라고도불림 캔형연소기의가장큰장점은가스터빈전체공기유량과연료유량가운데일부분만을이용하여연소기를개발할수있다는점임. 가스터빈개발은주로실험적연구를통해장기간수행되기때문에새로운연소기개발에적은공기유량과연료유량을이용하는것은개발비용측면에서큰장점 캔형연소기는가스터빈으로부터독립적으로분해가가능하기때문에유지보수가상대적으로쉬움 그러나캔형연소기는연소기출구에이웃한연소기와연결되는부위에배기가스가나오지않는비활동영역 (inactive area) 이만들어지기때문에연소기출구에서원주방향을따라배기가스유동분포가고르지못하며, 이로인해가스터빈진동이증가하는단점보유 6. Gas Turbines 64 / 140

65 환형연소기 [1/2] 환형연소기는연소실형상이원주방향을따라막힘없는환형 환형연소기는캔형연소기와달리하나의라이너와하나의케이싱으로구성 장점 동일한출력과직경을가지는캔형연소기에비해 75% 정도의길이를가지기때문에가스터빈무게와가격이크기줄어듬 캔형연소기에비해라이너표면적이작기때문에더적은냉각공기를필요로하며, 이로인해가스터빈열효율과출력이향상. 이는라이너표면적이작아지면연소기에서발생하는압력손실크기가줄어들뿐만아니라연소에더많은압축공기를사용할수있기때문임 화염전파에따른문제가없음 연소기출구에서원주방향온도분포및속도분포가캔형연소기에비해균일하기때문에가스터빈진동이줄어들고터빈블레이드수명이길어짐 단점 다수의연료노즐을사용하더라도연소기출구에서균일한온도분포를얻기힘듦 캔형연소기에비해구조적으로취약하기때문에고온상태에서라이너에좌굴이발생할가능성이높음 환형연소기를개발하기위해서는가스터빈전출력에해당하는질량유량의압축공기를공급하기위한방대한실험설비구축이요구되며, 많은연료가필요하기때문에과도한개발비용소요 분해하기어렵기때문에유지보수가어려움 6. Gas Turbines 65 / 140

66 환형연소기 [2/2] V94.3 & V84.3 [Siemens] V94.3 gas turbine consists of 16-stage axial flow compressor followed by an annular combustor and a four-stage reaction type axial-flow turbine. Annular combustors are superior to can combustors in terms of overall temperature distribution factor (OTDF). Can combustors have a relative higher OTDF that may result in thermo-mechanical fatigue problems. Annular combustor popularity increases with higher temperatures or low-btu gases, because the amount of cooling air required is much less than in can type designs due to a much smaller surface area. The amount of cooling air required becomes an important consideration in low-btu gas applications, because most of air is used up in the primary zone and little is left for film cooling. 6. Gas Turbines 66 / 140

67 Diesel engine Spark ignition engine Design Change of Combustors Diffusion combustor Wet combustor DLN combustor Steam or water injection Catalytic combustor Inclusion of catalyst Single fuel nozzle Multiple fuel nozzle Reduced NO x emission Low NO x emission Near zero NO x emission Premix fuel and air before combustion Fuel injector Spark plug 6. Gas Turbines 67 / 140

68 Different Modes of Laminar Combustion Diffusion Flame Fuel and air mix and burn at the same time An example for diffusion combustion is a Diesel engine, where a liquid fuel spray is injected into the compressed hot air within the cylinder. It rapidly evaporates and mixes with the air and then autoignition under partly premixed conditions Flame color is bright yellow NO x formation in post-flame regions Premixed Flame Fuel and air mixed and then burn In a spark ignition engine, a premixed turbulent flame front propagates from the spark through the combustion chamber until the entire mixture is burnt. Flame color is blue to bluish-green Low NO x burners Surface of stoichiometric mixture Post flame radiation Post flame oxidation and radiation Premixed flame front Air Air Fuel Fuel + air 6. Gas Turbines 68 / 140

69 DLN Combustor [1/2] The high costs of both water injection and SCR systems give opportunities to develop advanced combustors, so-called dry low NO x (DLN) combustors. Moreover, the introduction of steam or water to the gas turbine combustor is a thermodynamic loss, due to taking some of the energy from combustion gases to heat water or steam. However, DLN combustor has no impact on the cycle efficiency. Therefore, DLN combustor is more desirable than steam/water injection. DLN combustor premixes air and fuel, and makes a fuel lean mixture that significantly reduces peak flame temperature and thermal NO x formation. Another important advantage of the DLN combustor is that the amount of NO x formed does not increase with residence time. Since long residence times are required to minimize CO and unburned hydrocarbon (UHC) emissions, DLN systems can achieve low CO and UHC emissions while maintaining low NO x levels. To minimize flame temperature and hence NO x formation the fuel/air mixture is weakened to as near the extinction point as can safely be realized. The main problems associated with lean premix flames are stability, inflexibility and the limited turn-down range. To stabilize the flame, hybrid system having two fuel injectors of main fuel and pilot fuel is used commonly. In the hybrid system, the bulk of the fuel (more than 75%) is burned in a premixed burner, the remainder being supplied to a small pilot diffusion flame embedded in the flow. The main fuel is injected into the air stream immediately downstream of the swirler at the inlet to the premixing chamber. The pilot fuel is injected directly into the combustion chamber with little if any premixing. 6. Gas Turbines 69 / 140

70 DLN Combustor [2/2] A small portion of the fuel is always burned richer to provide a stable piloting zone, while the remainder is burned lean. In both cases, a swirler is used to create the required flow conditions in the combustor to stabilize the flame. With the flame temperature being much closer to the lean limit than in a diffusion combustor, some action has to be taken when the engine load is reduced to prevent flame out. If no action was taken, flame out would occur since the mixture strength would become too lean to burn. Due to flame instability limitations of the DLN combustor below approximately 50% of rated load, the combustor is typically operated in a conventional diffusion flame mode, resulting in higher NO x levels. DLN fuel injector is much larger because it contains the fuel/air premixing chamber and the quantity of air being mixed is large, approximately 50-60% of the combustion air flow. The operation is limited to a narrow range of fuel/air ratio between the production of excessive NO x and excessive CO. Some manufacturers are now offering dual-fuel DLN combustors. However, DLN operation on liquid fuels has been problematic due to issues involving liquid evaporation and auto-ignition. This consideration becomes more important as power producers consider converting from natural gas only to dual-fuel operation as natural gas price rise. 6. Gas Turbines 70 / 140

71 NO x and CO emissions, ppm Emissions in a DLN Combustion [1/2] 1) 순수한예혼합연소는잉여공기비 (excess air ratio) 가증가함에따라 NO x 발생이급격히줄어들며, 잉여공기비가 2 이하인경우 CO 발생은매우적다. 2) 순수한예혼합연소는운전영역이매우좁은데, 이는잉여공기비가 2 에가까워지면화염이꺼지기때문이다. 3) 예혼합연소에파일럿화염을적용하면화염이꺼지는염려가없이넓은잉여공기비에걸쳐서운전이가능하다. 이와같은이유때문에 GE 사와 Siemens 사모두파일럿화염을적용한예혼합연소기를개발하여사용하고있다. 4) 파일럿화염을적용한예혼합연소는 NO x 와 CO 의발생을최소화하기위해서운전영역을 NO x 와 CO 발생교차지점으로제한한다. 5) 확산연소는잉여공기비전영역에걸쳐서예혼합연소에비해 NO x 발생량이많다. 6) 확산연소의운전영역은예혼합연소에비해더큰잉여공기비를가지는부분에서형성되며, NO x 와 CO 의배출량도훨씬많다. Excess air ratio CCPP includes gas turbines with DLN combustors that can operate with stack gas NO x emission concentration as low as 25 ppmvd at 15% oxygen without steam or water injection, when the natural gas is used as a fuel. NO x can be reduced to less than 9 ppmvd by the installation of SCR in the HRSG. 6. Gas Turbines 71 / 140

72 NO x emission, ppmvd Flame temperature Emissions in a DLN Combustion [2/2] Reduction of emissions in the premix combustor NO x are reduced by Lowering flame temperature by lean combustion Elimination of local hot spots CO and UHC are reduced by Increasing combustion residence time (volume) Combustor design to prevent local quenching 250 Diffusion combustor Premixed combustor Extinction of lean premix flame Fuel lean Fuel rich Diffusion combustor DLN combustor Fuel: natural gas Fuel to air ratio Catalytic combustor 6. Gas Turbines 72 / 140

73 Diffusion vs. Premix Diffusion combustor Lean premix combustor Discuss the advantages of a lean premix combustor 1) Lower NO x emission Low flame temperature 2) Larger power output Less cooling air is required 3) Lower CO and UHC emission Increased residence time 4) Extended hot gas parts No water/steam injection 6. Gas Turbines 73 / 140

74 GE DLN Combustor [1/4] DLN-1 Combustor 1) The fuel-air equivalence ratio and residence time in the flame zone to be low enough to achieve low NO x. 2) Acceptable levels of combustion noise (dynamics). 3) Stability at part-load operation. 4) Sufficient residence time for CO burn-out. 6. Gas Turbines 74 / 140

75 Operating Modes of DLN-1 Combustor GE DLN Combustor [2/4] 6. Gas Turbines 75 / 140

76 Operating Modes of DLN-1 Combustor GE DLN Combustor [3/4] A small portion of the fuel is always burned richer to provide a stable piloting zone, while the remainder is burned lean. Primary Flame is in the primary stage only. This mode is used to ignite, accelerate and operate the machine over low- to mid-loads, up to pre-selected combustion reference temperature. Lean-Lean Flame is in both the primary and secondary stages. This mode is used for intermediate loads between two pre-selected combustion reference temperature. Secondary Flame is in the secondary stage only. This mode is a transition state between lean-lean and premix modes. This mode is necessary to extinguish the flame in the primary zone, before fuel is reintroduced into the primary zone. Premix Fuel to both primary and secondary zones. Flame is in the secondary stage only. Optimum emissions are generated in this mode by premixed flow. In the premix mode, the first stage thoroughly mixes the fuel and air and delivers a uniform, lean, and unburned fuel/air mixture to the second stage. A pilot nozzle produces a stable diffusion flame that can maintain high flammability in the premixed flame. 6. Gas Turbines 76 / 140

77 CO (ppmvd) NO 15% O 2 (ppmvd) GE DLN Combustor [4/4] Emission Level - GE DLN-1 Combustor (Fuel: NG) ISO Ambient Conditions NO x 10 CO Gas Turbine Load, % 6. Gas Turbines 77 / 140

78 Fundamentals of Gas Turbines Compressor Combustor Turbine 6. Gas Turbines 78 / 140

79 Turbine Turbine Air Inlet Compressor Combustors Exhaust Cold Section Hot Section The function of the turbine is to extract energy from the working fluid and convert it to mechanical energy, thereby enabling the turbine to drive the compressor and generator. 6. Gas Turbines 79 / 140

80 Configuration 6. Gas Turbines 80 / 140

81 터빈일반사항 터빈은고온의배기가스에포함된압력에너지와열에너지일부를기계적인일로변환시켜압축기와발전기를구동하기위한동력생산 압축기가공기에에너지를부여하여공기의압력을증가시키는반면에터빈은배기가스의압력을낮추어서배기가스에포함되어있는에너지를뽑아냄. 이는터빈을구성하고있는노즐과버켓에서배기가스가가속되면서배기가스의압력에너지가운동에너지로바뀌면서일어남 이렇게증가된배기가스의운동에너지는버켓에흡수되어기계적인일로변환. 따라서고속의배기가스는버켓을통과하는동안운동에너지를잃기때문에속도가크게줄어든상태로버켓을빠져나옴 즉터빈을통과하면서배기가스는팽창되어압력과온도가낮아짐 노즐과버켓은가장효율적으로에너지를뽑아내기위해공기역학기술을바탕으로설계 터빈은연소기바로하류에위치하고있기때문에터빈블레이드 ( 노즐과버켓 ) 는압축기블레이드 ( 로터와스테이터 ) 보다훨씬열악한환경에서운전 가스터빈엔진은높은열효율을얻기위해높은 TIT 요구. 따라서터빈블레이드는오랜운전기간동안블레이드의용융점을초과하는온도를견디어야함 이를위해터빈블레이드는효율적인냉각필요 아울러버켓은시뻘겋게달아오른상태에서고속회전에따른원심부하를견딜수있을만큼튼튼해야함 노즐과버켓은니켈합금을이용하여제작. 니켈합금은고온에서우수한크립 (creep) 특성보유 6. Gas Turbines 81 / 140

82 터빈동력생산원리 [1/6] 유체역학적힘 F = mv = V 2 A m = VA (mass flow rate) Nozzle A, V F R Reaction Action 6. Gas Turbines 82 / 140

83 터빈동력생산원리 [2/6] 터빈블레이드명칭 Leading Edge Blade Thickness Camber Angle Suction Side Pressure Side Trailing Edge Deflection Stagger Angle Pitch Blade Inlet Angle Blade Outlet Angle Gas Inlet Angle Gas Outlet Angle Direction of Gas Flow Incidence Tangential Deviation Angle Direction of Gas Flow Axial 6. Gas Turbines 83 / 140

84 유체유동에의해발생하는힘 터빈동력생산원리 [3/6] 1 V 1 m V 1 Tangential m V 1 sin1 V2 sin 2 Axial V 2 m V Gas Turbines 84 / 140

85 터빈동력생산원리 [4/6] 유체유동에의해버켓에발생하는힘의크기 배기가스는피치에해당하는면적에경사진형태로버켓통로로유입 따라서유동조건과버켓열이형성하는기하학적데이터를이용하면유입되는배기가스에의해버켓에접선방향으로작용하는힘의크기계산가능 이와같은방법으로버켓을빠져나가는유동조건을이용하면버켓을빠져나가는배기가스의반작용에의해발생하는접선방향힘의크기계산 그리고유입되는배기가스와배출되는배기가스에의해접선방향으로작용하는두힘의크기를합치면버켓에접선방향으로작용하는전체힘의크기가됨 그러나이방법으로는버켓에작용하는힘의크기를정확하게계산하기어려움. 그이유는버켓날개표면에서발생하는경계층때문에버켓을빠져나오는유동이균일하지못하기때문임 버켓에작용하는힘을계산하기위한또다른방법으로날개이론 이방법은버켓표면에작용하는압력분포를이용하여양력을계산하는방법으로써가장정확하면서실제적으로가장많이이용 흡입면압력이압력면에비해서낮으며, 이로인해버켓에양력발생 6. Gas Turbines 85 / 140

86 터빈동력생산원리 [5/6] 날개주위유체거동 NACA 4412 Velocity distribution p o 1 p1 V p 2 1 2V Pressure distribution 6. Gas Turbines 86 / 140

87 터빈동력생산원리 [6/6] 버켓단면에나타나는공기역학적현상을살펴보면, 배기가스가버켓을지나면서압력면 (pressure surface) 에흡입면 (suction surface) 보다높은압력형성 이로인해버켓압력면에서흡입면방향으로, 즉접선방향으로버켓을들어올리는양력발생 그런데버켓은터빈디스크에체결되어있기때문에버켓에발생하는양력은터빈축을회전시키는토크로작용하며, 이토크가압축기와발전기구동에사용되는회전력으로작용 버켓에서생산된양력에버켓이회전한거리를곱하면버켓이한일의크기가되며, 이일의크기가버켓에서생산된기계적인일의크기가됨. 한편, 일을시간으로나누면동력이됨 1 p 2 p 1 p c 1 ½ c 1 2 b Direction of Rotation P S S P c 2 2 P: Pressure Surface S: Suction Surface p 2 ½ c 2 2 p o 6. Gas Turbines 87 / 140

88 터빈구성품 1. 노즐 노즐은고정블레이드 (stationary blade) 로서흔히고정베인 (stator vane) 이라고도불리며, 버켓앞열에위치 노즐이조립되어있는것을노즐다이어프램 (nozzle diaphragm) 이라함 노즐의일차적인역할은연소기에서배출된고온가스에포함되어있는압력에너지의일부를운동에너지로변환시킴 노즐단면은일반적인비행기날개형상을가지고있으며, 이웃한노즐과의사이에형성되는노즐통로는수축형상을가짐 따라서고온가스가노즐을통과하는동안유동가속이일어나면서고온가스에포함되어있는압력에너지의일부가운동에너지로변환 노즐출구에서나타나는배기가스속도는가스터빈엔진전체에서나타나는속도가운데가장빠른속도 노즐의이차적인역할은노즐에서가속된배기가스를터빈버켓에최적의각도로유입되도록안내해서버켓이최대의효율을발휘하도록해줌. 버켓은배기가스입사각변화에따라성능과진동특성에큰영향을받음 노즐은가볍게하기위해대개가운데가비어있는형태이며, 이빈공간으로압축기로부터전달된공기를통과시켜냉각시킴 노즐 ( 고정블레이드 ) c 1 c 2 노즐열 노즐통로노즐노즐다이어프램 6. Gas Turbines 88 / 140

89 터빈구성품 1. 노즐 Nozzle 6. Gas Turbines 89 / 140

90 터빈구성품 2. 버켓 버켓을흔히블레이드, 또는회전블레이드 (rotor blade, or rotating blade) 라고도함 버켓은노즐에서가속된고속의배기가스운동에너지를압축기와보기류를구동하기위한기계적인일로변환시키는역할수행 크립 (creep) 특성이우수한니켈합금으로제작 노즐한열 (row) 과버켓한열을합쳐서한단 (stage) 이라함 6. Gas Turbines 90 / 140

91 터빈구성품 3. 슈라우드 슈라우드 : 버켓팁에달려있는일체형커버 슈라우드가설치되어있는블레이드를슈라우드블레이드 (shrouded blade) 라함 이에반해슈라우드가없는블레이드를자유팁블레이드 (free tip blade), 또는오픈팁블레이드 (open tip blade) 라함 슈라우드설치목적 : 1) 버켓팁을넘어가는배기가스양최소화를통한터빈효율향상 슈라우드는인접한버켓의슈라우드와연결되어버켓팁에연속적인링을형성하며, 그평평한바탕에나이프에지실 (knife-edge seal) 을설치하여배기가스누설을줄여가스터빈효율을향상시킴 슈라우드브레이드 버켓팁을넘어가는배기가스는터빈에서동력생산에기여하지못하기때문에손실이되며, 누설손실이증가하면가스터빈엔진효율저하 버켓팁으로과도한배기가스누설발생시버켓팁부위의효율을떨어뜨리는난류발생하며, 이로인해터빈효율감소 2) 버켓진동특성향상 슈라우드는이웃한버켓의슈라우드와서로맞물려버켓진동을감쇄시킴 슈라우드블레이드는자유팁블레이드에비해진동특성이우수한진동모드를가짐 슈라우드는이웃한슈라우드에맞물려버켓의뒤틀림 (distortion) 을방지함으로써얇고긴형상의가벼운버켓제작을가능케해줌. 이로인해슈라우드블레이드는가스터빈무게절감에도기여 슈라우드버켓을사용하면슈라우드로인해버켓단면에작용하는응력이증가하는경향을가짐 자유팁브레이드 6. Gas Turbines 91 / 140

92 터빈구성품 3. 슈라우드 3) 버켓팁부위공력특성향상을통한터빈효율향상 자유팁블레이드는팁에서팁와류 (tip vortex) 발생 슈라우드를장착하면팁와류발생방지하여터빈효율향상 6. Gas Turbines 92 / 140

93 터빈구성품 3. 슈라우드 일반적으로슈라우드블레이드와자유팁블레이드는하나의엔진에서동시에볼수있음 대부분의엔진에서고속으로회전하는블레이드는자유팁으로하며, 저속으로회전하는블레이드는슈라우드장착 따라서고압터빈은자유팁블레이드, 저압터빈은슈라우드블레이드사용 공기냉각을실시하는블레이드의경우자유팁으로하는것이일반적임 대부분의고압터빈버켓은냉각이요구됨 6. Gas Turbines 93 / 140

94 터빈구성품 4. 도브테일 버켓을터빈로터디스크에체결 도브테일은버켓에서발생하는원심응력을감당하기때문에매우높은가공정밀도요구 전나무형도브테일 (fir tree dovetail) 이많이사용 Shank Shank prevents heat transfer from turbine blade to dovetail. Vibration control. Reduce stress. Shank Dovetail 6. Gas Turbines 94 / 140

95 터빈열역학적고찰 [1/2] Air 1 2 Fuel Combustor 3 Exhaust gas 4 Power T (h) Compressor Turbine 1 p T h c o o o p T 터보팬엔진과터보프롭엔진개념도및 h-s 선도 = total pressure = total temperature = specific stagnation enthalpy = specific heat = specific heat ratio = isentropic efficiency of turbine TPR p p o,4 o,3 T T o,4 o,3 1 TW T T c T h p p h o, 3 o,4 T c T c T p T o, 3 o,4 T 1 T Tc o,4 o, 3 pto,3 1 o,3 T o, 3 1 TPR 1 p p o,4 o,3 1 s 6. Gas Turbines 95 / 140

96 터빈열역학적고찰 [2/2] 터빈에서일어나는과정을열역학적으로살펴보면, 즉터빈에서일어나는손실을무시하거나손실의크기가매우작다고가정하면, 터빈에서는등엔트로피팽창과정 ( 그림에서과정 34) 을통해동력이생산됨. 그러나실제적으로터빈에서는손실이발생하며, 발생하는손실을반영하면터빈에서는단열팽창과정이일어남 ( 그림에서과정 34) 즉터빈에서과정이진행되는동안터빈내부로열이공급되지도않으며, 터빈내부로부터외부로열이빠져나가지도않음. 그러나엄밀하게살펴보면, 터빈케이싱이가열되면서일부열이터빈케이싱을통해터빈외부로빠져나가기때문에단열과정아님. 그러나터빈을통과하는배기가스속도가너무빨라서미처열교환이일어날시간이없다고가정하면터빈에서단열과정이일어남 배기가스가터빈을통과하는동안전체압력과전체온도가낮아지는팽창과정이진행되며, 이팽창과정을위해압축기에서공기의압력을올리는것임 열역학제 1 법칙을이용하여터빈에서생산되는열역학적인일의크기를계산하면 (h o,3 -h o,4 ) 임. 여기서 h o 는정체엔탈피 (stagnation enthalpy). 엔탈피는내부에너지와유동에너지를합친값으로정의 그러나터빈에서발생하는다양한손실, 즉형상손실, 이차유동손실, 누설손실등으로인해실질적으로터빈에서생산되는일의크기는이상적인일의크기보다작아짐 터빈에서실질적으로생산되는일의크기는열역학적일의크기에터빈효율을곱한값 터빈에서일어나는배기가스압력강하는터빈압력비 (TPR; Turbine Pressure Ratio) 로나타냄. 터빈압력비는터빈출구압력 (p o,4 ) 을터빈입구압력 (p o,3 ) 으로나눈값으로정의 결과적으로터빈에서생산되는일의크기에관한관계식을살펴보면, 터빈에서생산되는일의크기는터빈압력비, TIT, 정압비열및비열비와같은공기의상태량, 그리고터빈효율에영향을받음 6. Gas Turbines 96 / 140

97 터빈유체역학적고찰 [1/7] 터빈단에서의유체거동 r z 노즐열 버켓열 6. Gas Turbines 97 / 140

98 터빈유체역학적고찰 [2/7] 터빈에서일어나는유동 (flow behavior) 은한단을기준으로조사 단입구 ( 노즐입구 ) 에서유동은반경방향을따라일정하게유지. 아울러단출구 ( 버켓출구 ) 에서의유동또한반경방향을따라일정하게유지. 따라서다단축류터빈 (multi-stage axial turbine) 설계용이 배기가스는노즐을통과하면서가속됨. 그러나대부분의축류형터빈설계에있어축방향속도성분크기는다음단설계의편의를위해단을통과하는동안일정하게유지되도록설정. 따라서노즐을빠져나온배기가스는큰접선방향속도성분가지며, 이로인해노즐출구와버켓입구사이에는큰선회유동 (swirl flow) 형성 그러나노즐출구환형면적에서나타나는선회유동을 3 차원으로나타내기어렵기때문에터빈에서나타나는유동을정확하게파악하기어려움 이런어려움을극복하기위하여일반적으로그림에나타나있는원통의옆면에나타나는유동을살펴보는방법사용 원통옆면을평면상에펼쳐놓으면 3 차원유동이 2 차원유동이되기때문에유동패턴쉽게분석가능 이경우접선방향 ( 원주방향 ) 을따라유동이반복되기때문에노즐과버켓모두서로인접한블레이드사이에형성되는하나의유동통로에서나타나는유동을살펴보면됨 그리고원통직경을버켓루트에서팁까지증가시키면서살펴보면터빈단에서나타나는전체적인유동을정확하게조사가능 이렇게터빈한단에대해조사한속도성분은속도삼각형 (velocity triangle) 으로나타남 6. Gas Turbines 98 / 140

99 터빈유체역학적고찰 [3/7] 절대속도 vs. 상대속도 절대속도 (absolute velocity) 정지상태에서관찰한속도 어떤물체의이동속도그자체 p 1 c 1 1 c : 절대속도 w : 상대속도 u : 버켓회전속도 속도삼각형에서 c 로표기 노즐열 상대속도 (relative velocity) 이동상태에서관찰한속도 상대속도 = 관찰대상물체의이동속도 관찰자속도 속도삼각형에서 w 로표기 버켓으로유입되는배기가스의속도는버켓이회전하고있기때문에상대속도관점에서살펴보아야함 p 버켓열 u w 2 c 2 u 3 버켓의접선방향회전속도가 u 이기때문에버켓으로유입되는배기가스의상대속도크기 w = c - u p 3 w 3 3 c 3 u 6. Gas Turbines 99 / 140

100 터빈유체역학적고찰 [4/7] 루트에서의속도삼각형 Tip p 1 c 1 1 Nozzle Row Root p w 2 c 2 u r Root r Tip Bucket Row u Nozzle Row Bucket Row 3 p 3 w 3 3 c3 c : absolute velocity of fluid u : tangential velocity of bucket w : relative velocity of fluid to bucket u 접선방향 축방향 6. Gas Turbines 100 / 140

101 터빈유체역학적고찰 [5/7] 팁에서의속도삼각형 Tip p 1 c 1 1 Nozzle Row Root p w 2 c 2 u r Root r Tip Bucket Row u Nozzle Row Bucket Row p 3 w 3 3 c3 3 u 6. Gas Turbines 101 / 140

102 터빈유체역학적고찰 [6/7] 노즐로유입된배기가스는노즐을통과하면서가속 (c 1 c 2 ) 노즐에서반동터빈은 2 배, 충동터빈은 4 배가속 노즐통과후에도축방향속도성분크기는변함없음. 따라서접선방향속도가크게증가함. 버켓은접선방향을따라회전하기때문에이증가한접선방향속도성분이버켓회전력생산에기여 버켓은회전하고있기때문에절대속도인 c 2 가버켓으로유입되는것이아니라상대속도인 w 2 가유입 버켓앞전 (leading edge) 은상대속도 w 2 에적합한각으로정렬됨 버켓을빠져나가는배기가스의절대속도 (c 3 ) 는노즐로유입되는배기가스절대속도 (c 1 ) 와크기및방향동일. 이것이다단축류터빈기본설계원리 그러므로버켓회전속도와버켓출구에서배기가스의절대속도가정해졌기때문에버켓출구에서의배기가스상대속도 w 3 가결정됨 따라서버켓뒷전 (trailing edge) 정렬각이결정되면서설계단계에서선정된버켓형상의캠버가결정되면서버켓형상이완성됨 이렇게터빈한단에서는노즐출구와버켓출구에서각각한개씩총두개의속도삼각형생성 두속도삼각형에서버켓으로유입된절대속도크기 (c 2 ) 와버켓을빠져나간절대속도크기 (c 3 ) 를비교해보면버켓을통과하는동안배기가스의절대속도크기가크게줄어든것을확인할수있음 즉버켓을통과하는동안배기가스의운동에너지가크게줄어들었다는것을알수있으며, 줄어든운동에너지는버켓에서터빈축을회전시키는기계적인일로변환 6. Gas Turbines 102 / 140

103 터빈유체역학적고찰 [7/7] 관찰면을루트에서팁으로이동시켜살펴본후루트에서의속도삼각형과비교하면버켓의입체적인형상을살펴볼수있음 편의상노즐을빠져나오는배기가스속도는서로동일하다고가정 팁에서버켓회전속도는루트에비해증가. 이는팁반경 (r Tip ) 이루트반경 (r Root ) 에비해크기때문임 따라서버켓으로유입되는배기가스상대속도가작아지며터빈축과배기가스상대속도가이루는각 ( 2 ) 크기가작아짐 버켓을빠져나가는배기가스의절대속도 (c 3 ) 는노즐로유입되는배기가스절대속도 (c 1 ) 와크기및방향동일. 그런데팁에서버켓회전속도가증가하기때문에버켓팁을빠져나가는배기가스속도 w 3 는루트에서의크기보다증가하며, 이로인해루트에비해노즐출구면적축소 아울러팁에서버켓을빠져나가는배기가스방향과터빈축이이루는각 ( 3 ) 크기가루트에비해증가 이런이유때문에버켓은루트에서팁으로비틀린형상을가짐 6. Gas Turbines 103 / 140

104 반동도 회전블레이드에서의엔탈피변화 = x 100 (%) 단에서의엔탈피변화 h h h h % T T T T % p p p p % q dh dp 버켓에서의엔탈피변화량을단에서의엔탈피변화량으로나눈값으로정의하는반동도 (degree of reaction) 는터빈설계특성을구분하는매우중요한무차원수 터빈을통과하는배기가스의엔탈피는측정하기어려움 엔탈피는정압비열과온도와의곱이므로반동도는온도의함수로표시가능 터빈을통과하는배기가스의온도역시실질적으로측정하기어려움 그러므로반동도관계식을터빈단에서측정이비교적쉬운열역학적상태량인압력으로표시가능 압축기와터빈에서일어나는열역학적과정은단열과정. 따라서 q = 0 성립 터빈단에서일어나는밀도변화가매우작다고가정하면, 열역학제 2 기초식으로부터 dh dp 성립 6. Gas Turbines 104 / 140

105 충동터빈 충동터빈 (impulse turbine): 반동도 0% 를가지는터빈단으로구성된터빈 터빈단에서압력강하는노즐에서만일어남 노즐열 버켓에서는압력강하가일어나지않기때문에버켓을통과하는동안배기가스속도는일정하게유지. 그러므로버켓통로단면적은버켓입구로부터출구에이르기까지일정. 이런조건을만족시키기위해충동터빈버켓단면은초승달형상임 버켓열 충동터빈은고속배기가스의운동방향을바꾸어서에너지흡수 V j U F V j U 버켓 [ 충동터빈, = 0% ] U 6. Gas Turbines 105 / 140

106 반동터빈 반동터빈 (reaction turbine): 반동도 50% 를가지는터빈단으로성된터빈 즉반동터빈은단에서의압력강하가노즐과버켓에서각각절반씩일어남. 그러므로노즐과버켓은동일형상을가짐 Nozzle Row 노즐과버켓단면은일반적인날개형상을가짐 서로인접한노즐과버켓사이에형성되는유동통로단면적은하류로가면서좁아지는형상을가짐. 그러므로배기가스가노즐과버켓유동통로를지나면서압력은낮아지고속도는증가 Bucket Row 버켓을고속으로빠져나가는배기가스의반력으로버켓은회전동력을얻음 V i F V e U Convergent nozzle [ 반동터빈, = 50% ] 6. Gas Turbines 106 / 140

107 반동터빈 Hero s Aeolipile (BC 150 년경 ) 6. Gas Turbines 107 / 140

108 터빈단에서의상태량변화비교 Impulse Turbine Reaction Turbine 노즐 버켓 노즐 버켓 c 1 c 1 w 2 w 3 w 2 w 2 w 3 A 1 w 2 A 2B A 3B A 1 c 2 u A 3B c 2 u A 2B w 3 A 2N 3 u c 1 A 2N 2 2 w 2 w 3 3 u c 1 2 c 3 p 1 c 2 u c 3 3 = 0 p 1 p 2 c 2 w 2 2 = 0 U 3 = 0 T 1 p 2 p 3 T 1 T 2 p 3 T 2 T 3 T 3 A 2N A 1 A 2B = A 3B 2 = 3 w 2 = w 3 c 2 4c 1 A 2N A 1 A 3B A 2B c 2 c 1 w 3 w 2 c 2 2c 1 6. Gas Turbines 108 / 140

109 충동 - 반동터빈 [1/3] 노즐의역할은작동유체의압력에너지를운동에너지로변환시키는것이다. 따라서노즐을통과한작동유체의속도는크게증가한다. 그러나노즐출구에서의축방향속도는노즐입구에서와동일하기때문에접선방향속도만크게증가한다. 이로인해노즐출구를빠져나온작동유체는큰선회유동으로인해원심력이발생하여유체는버켓팁 (tip) 쪽으로집중되는경향을가진다. 유동이버켓팁쪽으로편중되면버켓과케이싱사이에서누설손실이증가하며, 버켓팁근처에서이차유동손실이증가할뿐만아니라반경방향을따라서버켓에서생산하는동력도균일하지못하게된다. 이런문제를해결하기위해서버켓팁입구쪽의압력을루트 (root, or hub) 입구쪽압력보다높게유지시킨다. 이경우버켓입구의팁부분압력이루트부에비해높기때문에팁쪽에서루트방향으로진행하는유동이형성된다. 즉버켓팁쪽에형성되는높은압력으로인해루트쪽으로진행하려는힘과원심력에의해루트에서팁쪽으로진행하려는두힘은서로방향이반대이기때문에두힘의크기를비슷하게해주면노즐과버켓사이에서유동은축방향으로평행하게흘러가며, 앞서언급된제반문제점들이사라지게된다. 따라서노즐과버켓사이에형성되는유동의특징은반경방향을따라서속도는줄어들고, 압력은증가한다. 한편, 축류형다단터빈은버켓입구에서뿐만이아니라출구에서도압력과속도는반경방향을따라서일정하게유지되어야한다. 따라서버켓은루트에서팁쪽으로가면서반동도가증가하기때문에버켓루트는충동형, 팁은반동형으로설계한다. 이런이유때문에터빈버켓은하나의블레이드에충동형과반동형이혼재된충동 - 반동블레이드 (impulsereaction blade) 이다. 6. Gas Turbines 109 / 140

110 버켓팁으로의유동편중해결방법 충동 - 반동터빈 [2/3] Free vortex design impulse type HP turbine stage 1000 psia psia tip psia 1000 psia psia pitch psia 1000 psia psia root psia [ Example of pressure variation in radial direction ] 6. Gas Turbines 110 / 140

111 Radial Variation of Flow Parameters 충동 - 반동터빈 [3/3] w 2T c 2T u T c 1 c 3 c 2 p 1 p 2 p 3 w 2M c 2M u M c 2R w 2R u R 6. Gas Turbines 111 / 140

112 Stress 크립 [1/4] 각각의재료는모두고유의응력변형도곡선보유 그림에상온에서의연성재료인장시험에대한가장대표적인응력변형도곡선표시 c 연성재료에대한응력변형도곡선은크게두개의영역, 즉탄성변형영역과소성변형영역으로구분 탄성변형영역은그림에나타나있는구간 o-a 로써재료에가하는힘을제거하면재료는원래상태로돌아감. a b a: yield stress c: ultimate stress d: fracture o-a: elastic behavior a-d: plastic behavior d 그러나재료에항복응력을초과하는힘이가해지면힘을제거하더라도소성변형이발생하기때문에재료는원래상태로돌아가지못함. 즉재료에항복점보다큰응력이작용하면재료가잡아당기는힘을견디지못하며, 이로인해재료를구성하고있는물질의이동이일어나면서소성변형발생 o e Permanent Elongation Strain 따라서만약그림에나타나있는상태 b 에서재료에가해지는힘을제거하면상태 o 로복귀하지못하고상태 e 로복귀하기때문에재료에 o-e 만큼의영구신장발생 6. Gas Turbines 112 / 140

113 크립 [2/4] 가스터빈운전중버켓에는고속회전에따른원심력에의해인장응력이발생하며, 항복점보다훨씬작은응력이작용하더라도한엔진사이클 ( 가열 - 운전 - 정지 ) 경과후에는길이가약간늘어남 고온환경에서대부분의연성재료는항복점보다훨씬작은응력이작용하더라도응력을제거하면재료에약간의영구신장이일어나는데, 이를크립 (creep) 이라함 그러나정상적인운전상태에서크립에의한영구신장길이는엔진사이클당수백만분의 1 인치. 그러나온도가비정상적으로높아지거나과속상태에서운전하는경우정상운전에비해훨씬긴길이의영구신장발생 정상운전상태에서라도가스터빈을장기간운전하면크립에의한영구신장길이가누적되어버켓길이가늘어남. 이런경우늘어난버켓과케이싱사이에접촉이일어나서마찰 (rubbing) 이발생하며, 이로인해버켓팁일부가갈려나가면서심한마찰음발생 아울러버켓중간부위에넥킹이발생하면서버켓날개단면적이축소되는동시에버켓비틀림풀림. 이런경우버켓의공기역학적효율이크게낮아질수있으며, 심한경우넥킹이일어난부분이파단되어대형사고로이어질수있음 따라서가스터빈버켓은우수한크립특성을보유한소재를이용하여제작해야함 6. Gas Turbines 113 / 140

114 Strain 크립 [3/4] 일반적으로크립은 3 단계로구분 1 단계크립은가스터빈최초기동시일어나며, 빠르게진행됨 이에반해 2 단계크립은매우느리게진행 가스터빈제작자는가스터빈의운전수명을 2 단계크립이내로설정 3단계크립은급격히가속되는경향을가짐크립가속원인 Primary Phase Secondary Phase 설계온도보다높은온도에서운전 고출력상태에서장기간운전 외부물질유입에의한버켓침식 Tertiary Phase Time 6. Gas Turbines 114 / 140

115 크립 [4/4] 크립이진행되면입자경계 (grain boundary) 에크랙이발생하기시작 하중이작용하는방향을가로지르는방향의입자경계에서주로크랙이발생 Schematic of material placed in tension with a small elastic extension The effect of extended service on material structure at an elevated temperature, with the material subject to a tensile stress 6. Gas Turbines 115 / 140

116 터빈블레이드소재 [1/4] 초합금 (superalloys) 은다양한임계금속 (critical metals), 즉니켈, 크롬, 코발트, 티타늄, 텅스텐, 카본, 그리고다른금속요소의혼합체로구성 그러나가스터빈부품제작자들사이에이들임계금속의정확한혼합비는정해져있지않으며, 아직도큰논쟁거리로남아있음. 이에대한한가지이유는가스터빈부품의강도성질이궁극적으로이들금속요소의혼합비에좌우되기때문 그리고가스터빈부품소재의강도가증가할수록복잡한형상을가지는가스터빈부품으로성형하거나가공하기어렵기때문에비싸짐 초합금은고온환경에서산화저항성이요구되면서높은열응력, 인장응력, 진동에의한응력이작용하는부품에사용하기위하여개발 초합금이견딜수있는최대배기가스온도는냉각이수반되지않는버켓은 900 C, 냉각이수반되는버켓은 1,400~1,600 C 인코넬이라고도불리는니켈초합금 (nickel-base superalloys) 은철성분이포함되어있지않거나포함되어있더라도극소량이기때문에비부식성이며, 용접가능한얇은판재로만들수있음. 따라서니켈초합금은가스터빈연소기라이너, 터빈케이싱, 터빈블레이드등의제작에사용 가스터빈블레이드는새로운분말야금기술을이용한단조, 또는전통적인방법, 또는니켈초합금을이용한정밀주조방법으로제작. 이가운데정밀주조방법을이용하는경우응고속도및방법에따라서소재를구성하는입자의구조와소재의기계적특성이달라지기때문에등축결정, 일방향응고, 단결정세가지로분류 6. Gas Turbines 116 / 140

117 터빈블레이드소재 [2/4] 앞서설명했듯이크립이진행되면서입자경계에크랙이발생하며, 크랙은주로응력이작용하는방향을가로지르는방향의입자경계에서발생 등축결정소재는모든축을따라서소재의입자구조를균일하게배치해서크립을최소화시킨소재 미국의 P&W 사가 1965 년개발한일방향응고는입자경계를응력이작용하는방향과평행하게배치하고응력이작용하는방향을가로지르는방향의입자경계를없앤소재. 일방향응고소재는등축결정소재에비해크립강도, 열피로저항성, 부식저항성등이훨씬우수 비록일방향응고소재의입자경계가하중이작용하는방향으로평행하게형성되어있다고하더라도입자경계는기본적으로약하게연결되어있기때문에이들입자경계를강하게하려는많은기술들이개발 Low Creep characteristics High Ingot Price: Application in aviation: Application in power gen: 그러나입자경계를강하게하는것보다좋은것은입자경계를완전히없애는것임단결정소재는모든입자경계를없앰으로써소재의용융점을높여서고온강도를높인합금 단결정소재는일방향응고소재보다크립강도, 열피로저항성, 부식저항성모두우수 단결정소재는등축결정및일방향응고소재에비해저주기피로수명이 10% 더연장 그러나단결정소재는일방향응고소재보다훨씬고가 6. Gas Turbines 117 / 140

118 Relative Life Elongation 터빈블레이드소재 [3/4] 10x 8x Equiaxed DS SC Single Crystal Blades Directional Solidification Blades Fracture 6x Equiaxed Blades 4x 2x 0 Creep strength Thermal fatigue resistance Corrosion resistance Time 6. Gas Turbines 118 / 140

119 터빈블레이드소재 [4/4] 노즐은정지상태로존재하기때문에버켓과는달리원심력에의한인장응력을견딜필요없음 따라서노즐에요구되는가장중요한특성은내열성 노즐소재로는고온에서의용융을방지하기위해냉각이요구되기는하지만내열특성이우수한니켈합금사용 노즐에세라믹코팅을실시하면열저항특성이더욱향상되며, 냉각공기양을줄일수있기때문에가스터빈엔진효율향상 터빈버켓은고속회전으로발생하는높은원심하중을견딜수있을만큼튼튼해야함 비록무게가 60 그램인소형버켓일지라도최고회전속도에서 2 톤을초과하는부하가작용할수있음. 그리고압축기를구동하기위해수천마력을생산하기위한배기가스에의해발생하는높은굽힘응력을견디어야함 터빈버켓은높은주파수로변동하는배기가스의열특성에의한영향을받지않기위해피로 (fatigue) 및열충격 (thermal shock) 에대한저항성을가지고있어야함 터빈버켓은이런모든요구조건을만족하면서현재개발되어있는제작방법으로정확하게성형하고가공할수있는소재로제작되어야함 6. Gas Turbines 119 / 140

120 Cooling Methods Squealer tip Tip cap cooling holes Film cooling holes Film cooling Hot gas Rib turbulators Shaped internal cooling passage Tip cap cooling Trailing edge ejection Hot gas Trailing edge cooling slots Blade platform cooling holes Dovetail Impingement cooling Rib turbulated cooling Pin-fin cooling Cooling air Cooling air Source: Recent Advances of Internal Cooling Techniques for Gas Turbine Airfoils Minking K Chyu and Sin Chien Siw, J. Thermal Sci. Eng. Appl. 5(2), (May 17, 2013) 6. Gas Turbines 120 / 140

121 Cooling Air Extraction Cooling Air Supply and Its Flow 6. Gas Turbines 121 / 140

122 터빈블레이드냉각일반사항 가스터빈효율은 TIT 가증가할수록향상되기때문에 TIT 를증가시키기위한기술이지속적으로개발되고있음. 이가운데가장대표적인것이고온에견딜수있는블레이드 ( 노즐과버켓 ) 제작에사용하기위한내열소재개발과냉각기술개발 터빈블레이드는모재로사용하는니켈합금의용융온도보다약 500~700C 높은온도에서운전. 이는노즐과버켓을여러가지방법으로냉각시킴으로써가능 가장일반적인터빈블레이드냉각방법은압축기에서냉각공기를공급받아냉각시키는것임. 그런데압축기로부터공급되는냉각공기온도는어떤경우 600C 를초과할정도로고온임. 그러나이온도는터빈핵심부품인블레이드모재의용융온도보다낮음 냉각공기가터빈블레이드내부에형성된냉각공기통로를통과하면서블레이드열의일부를흡수하여블레이드냉각. 이를대류냉각 (convection cooling) 이라함 터빈블레이드내부를통과한냉각공기가블레이드표면에있는무수히많은작은구멍을통해빠져나와블레이드표면에차가운냉각공기막 (cooling air film, or protective blanket) 을형성하여뜨거운배기가스가블레이드표면에접촉하는것을막아줌으로써블레이드표면온도상승방지. 이를막냉각 (film cooling) 이라함 최근들어고온이형성되는블레이드앞전부위에냉각공기를고속으로충돌시켜냉각시키는기법이개발되어적용. 이를충돌냉각 (impingement cooling) 이라함 TIT 를증가시키기위하여블레이드표면을세라믹으로코팅하는기술이적용. 이를열차폐코팅 (TBC; Thermal Barrier Coating) 이라함 세라믹과같은코팅재료는블레이드모재보다더높은온도를견딜수있으며, 얇은코팅층은절연층 (insulating layer) 역할수행 세라믹과같은코팅재료는열전도도가낮기때문에블레이드에코팅층을형성시키면고온의연소가스로부터블레이드로전달되는열전달양을줄임으로써모재의표면온도를 100~150C 낮추어줄수있음. 즉 TIT 를블레이드표면온도를낮춘크기만큼올릴수있음 6. Gas Turbines 122 / 140

123 Temperature, C 터빈블레이드소재및냉각기술역사 Turbine inlet temperature Film cooling TBC S-816 N80A M252U500 Benefits of cooling U700 IN738 IN939 IN792 DS Single crystal Year Material improvement 6. Gas Turbines 123 / 140

124 1. 대류냉각 [1/2] 열전달메커니즘가운데대류열전달을이용한냉각방법. 현대식가스터빈에가장널리사용되는냉각방법 냉각공기가노즐과버켓내부를흘러가면서냉각통로벽면의열을흡수함으로써냉각 냉각면적으로증가시키기위해초기의단일통로 (single pass) 에서구불구불 (serpentine passage) 한다중통로 (multi-pass) 로개선 열전달계수를증가시키기위해냉각공기통로표면에터뷰레터 (turbulator; 난류발생장치 ) 설치 터뷰레터는열전달면적을증가시키는역할도수행 q ha T w T Film cooling Hot gas Impingement cooling Rib turbulators Shaped internal cooling passage Tip cap cooling Trailing edge ejection Rib turbulated cooling Pin-fin cooling Cooling air 6. Gas Turbines 124 / 140

125 1. 대류냉각 [2/2] Pin-Fin Cooling 가스터빈블레이드뒷전부분은매우얇기때문에복잡한냉각공기통로를설치하기어렵다. 따라서이부분에는핀 - 휜을설치하여효과적인냉각을도모하고있다. 핀 - 휜사이로지나가는냉각공기는핀이설치되어있는양쪽벽면과핀을통해전달되는열을흡수함으로써블레이드뒷전을냉각시킨다. 각각의핀주위에는하나의실린더주위에형성되는유동과거의유사한형태의유동이형성된다. 즉그림에나타나있는것처럼냉각공기가핀을통과하면서발생하는유동박리에의해후류가형성되어핀하류로흘러간다. 아울러핀바로앞부분에말발굽와류가발생하고핀을감싸면서흘러간다. 후류와말발굽와류모두난류형성을촉진하고냉각공기혼합을촉진하기때문에냉각효율향상에기여한다. 일반적으로많이사용되는핀 - 휜배열형태로는일렬형과엇갈림형이있다. Pin-fin 6. Gas Turbines 125 / 140

126 2. 막냉각 [1/2] 압축기마지만단에서추출된냉각공기는블레이드냉각공기통로를따라흐르면서블레이드를냉각시킨후에블레이드표면에있는일련의작은구멍을통해분사되어블레이드표면과고온의연소가스인주유동사이에절연층형성 이절연층을형성하는냉각공기막은고온의연소가스가블레이드표면에직접접촉하지못하도록해줌 아울러이냉각공기막은열을흡수함으로써블레이드표면으로열이전달되는것을줄여줌 6. Gas Turbines 126 / 140

127 2. 막냉각 [2/2] 블레이드냉각에대한역사를살펴보면, 막냉각적용을통해 TIT 가획기적으로증가했다는것을알수있다. 그림에 Westinghouse 사의블레이드냉각변천사에서도알수있듯이막냉각은블레이드냉각에가장효과적인냉각방법가운데하나라것을알수있다. Temperature F W501A (1968) And W501AA (1970) W501B (1973) Turbine inlet 1630 Average metal 1613 Maximum surface 1700 Cooling T 17 Turbine inlet 1819 Average metal 1452 Maximum surface 1650 Cooling T 357 W501D5 (1980) Turbine inlet 2025 Average metal 1390 Maximum surface 1615 Cooling T F (1990) Turbine inlet 2300 Average metal 1400 Maximum surface 1600 Cooling T Gas Turbines 127 / 140

128 3. 충돌냉각 충돌냉각 (impingement cooling) 은열부하가가장크게나타나는블레이드앞전에적용 특히블레이드앞전형상은대개뭉툭하기때문에충돌냉각을적용하기에적합 냉각공기가벽면과고속으로충돌하면열전달계수가증가되기때문에냉각효율이향상 뒷전냉각공기방출 핀 - 휜냉각 따라서냉각공기를고속으로블레이드안쪽면에충돌시키면더많은열이블레이드표면으로부터냉각공기로전달 일반적으로충돌냉각은블레이드전체표면에걸쳐서균일한온도분포를만들기위하여제한적으로사용 막냉각 충돌냉각 충돌냉각 배기가스 6. Gas Turbines 128 / 140

129 4. 침출냉각 침출냉각 (transpiration cooling) 은터빈블레이드를다공성재료로만들고블레이드내부에형성되어있는빈공간을통하여공급되는냉각공기가블레이드표면으로스며나오게함으로써터빈블레이드를냉각시키는방법 침출냉각은블레이드전체표면을고르게냉각시킬수있기때문에효과적인냉각방법 그러나다공성재료는크랙이발생하기쉽기때문에항상진동에노출되는가스터빈블레이드소재로적합하지못함 하지만기계적강도를지닌다공성내열소재가개발된다면침출냉각은가스터빈블레이드냉각에크게기여가능 Hot gas Cooling air film Porous material Cooling air 6. Gas Turbines 129 / 140

130 Temperature Nickel supperalloy substrate 5. 열차폐코팅 [1/2] Oxidation resistant bond coat: m Thermally grown oxide (TGO) Thermal barrier coating (TBC): m Interior cooling air Hot gases Hot gas temp. Thermal barrier coating C Cooling air temp. Distance 6. Gas Turbines 130 / 140

131 5. 열차폐코팅 [2/2] 일방향응고및단결정주조기술개발을통해가스터빈블레이드소재인니켈합금의운전온도가상승하였으며, 결과적으로이런소재기술의개발은 TIT 향상에기여 그후가스터빈소재기술개발은열차폐코팅 (TBC; Thermal Barrier Coating) 으로이동 열차폐코팅이란고온가스로부터터빈블레이드모재로전달되는열을줄여주기위해매우낮은열전도도를가진세라믹을이용하여블레이드표면에얇은두께로코팅함으로써절연층을만들어주는것임 이렇게블레이드표면에절연층이형성되어있으면블레이드표면을고온의배기가스로부터절연시켜블레이드표면온도를낮게유지시켜줄수있음 열차폐코팅은두개의코팅층, 즉본드코팅과탑코팅으로구성 본드코팅은블레이드의산화저항성과부식저항성을향상시키기위해실시하며, 본드코팅재질로는확산알루미나이드 (diffusion aluminide) 또는 MCrAlY 를사용 고온의배기가스열이블레이드로전달되는것을차단하기위해실시하는탑코팅재료는세라믹이며, 코팅층두께는 100~300 μm 정도이며, 모재온도를약 100~150 C 낮추어줌 코팅을실시하면블레이드수명이늘어나는데, 이는코팅층이손상을받으면코팅층을벗겨낸후코팅을다시실시하여블레이드를계속해서사용할수있기때문임 코팅층의수명은코팅재료조성, 두께, 균일도수준에의해좌우 터빈블레이드에세라믹과알루미늄합금열차폐코팅을실시하면모재의표면강도와부식저항성향상 따라서이런코팅은고온의배기가스유동에의해블레이드에발생하는비늘형태의부식 (scaling type corrosion) 또는침식을방지하는최선의방법 가스터빈블레이드는공기중에포함되어있는나트륨과연료에포함되어있는황 (sulfur) 이모재와화학적으로반응하면서부식이발생하기때문에부식에대한대비필요 6. Gas Turbines 131 / 140

132 Leading edge of blade Formation of Horseshoe Vortex Advanced Vortex Blade [1/8] Root Leakage (4%) Shaft Packing Leakage (7%) Tip Leakage (22%) Rotation (3%) Carryover (4%) Nozzle Profile (15%) Bucket Secondary (15%) Bucket Profile (15%) Nozzle Secondary (15%) Horseshoe vortex formed around a square bar s 1 s 2 Endwall Horseshoe vortex formed around a round bar 6. Gas Turbines 132 / 140

133 Formation of Horseshoe Vortex Advanced Vortex Blade [2/8] Endwall flow produces endwall boundary layer. Endwall flow is one of major sources of turbine losses, especially in cascades with short length blades and high flow turning. The endwall losses occupy a substantial part of the total aerodynamic losses in a nozzle or bucket row, even as high as 30~50%. Inlet boundary layer Stream surface Endwall The boundary layer fluid upstream of the leading edge is decelerated by the adverse pressure gradient and separates at a saddle point s 1. Passage vortex Counter vortex The boundary layer fluid elements form a reverse recirculating flow just before the leading edge. Endwall crossflow This reverse flow produces another saddle point s 2. The upstream boundary layer is rolled-up in the recirculating zone and it is divided into two legs at the leading edge saddle point of the blade and forms the so-called horseshoe vortex. Then, one leg goes into suction side and the other leg goes into pressure side in axial cascades. 6. Gas Turbines 133 / 140

134 Advanced Vortex Blade [3/8] Secondary Flow Secondary flow means various vortices passing through blade-to-blade passage in axial turbines. Stream surface Inlet boundary layer Endwall The pressure side leg moves towards the suction side of neighboring blade in the passage due to the tangential pressure gradient and becomes the passage vortex. Passage vortex Counter vortex The suction side leg called as the counter vortex rotates in the opposite direction to the larger passage vortex. There are two distinct (but arising from the same physical phenomena) vortices are present on the suction side of blades and they may merge, interact or stay separate. Counter vortex is also called as stagnation point vortex, or leading edge vortex, or horseshoe vortex. Endwall crossflow 6. Gas Turbines 134 / 140

135 Radial height Secondary Flow Loss Stream surface Advanced Vortex Blade [4/8] Inlet boundary layer Endwall Tip High efficiency area Passage vortex Counter vortex Hub Endwall crossflow Bucket efficiency Free Vortex Design Advanced Vortex Design 6. Gas Turbines 135 / 140

136 Concepts for an Advanced Vortex Blade Advanced Vortex Blade [5/8] Free vortex blade Leaned blade Compound leaned blade (Advanced vortex blade) 6. Gas Turbines 136 / 140

137 Evolution of Turbine Blades Advanced Vortex Blade [6/8] In order to reduce the secondary flow losses in turbine stages, radial velocity components are accounted for by using CFD techniques. Radial flow distribution is biased toward the more efficient mid-section of the bucket by the redistribution of the exit angle of nozzle blade. According to the open literature from GE, stage efficiency of the steam turbine can be improved 0.5 to 1.2% by the employment of advanced vortex blades. Nozzle solidity is reduced to allow use of more efficient blade profiles. Root reaction is moderately increased to increase efficiency, and tip reaction is decreased to reduce bucket tip leakage losses. [ M501G stage 3 vane segment ] [ M501G stage 4 vane segment ] 6. Gas Turbines 137 / 140

138 Efficiency Gain with 3D Blades Advanced Vortex Blade [7/8] Secondary flow loss can be reduced remarkably by the adoption of leaned blades Source: Siemens 6. Gas Turbines 138 / 140

139 Evolution of the First Stage Bucket Advanced Vortex Blade [8/8] The efficiency of an axialflow turbine, which in modern new advanced gas turbine, reaches about 92%. actual mean Tip Annulus height Increased inlet angle w u u u Reduced c Design c Increased c Inlet Middle ~ Exit Hub Decreased inlet angle 6. Gas Turbines 139 / 140