Off-grid test system for gas turbine engine and test method using the same

Abstract

The present invention relates to an off-grid test of a gas turbine engine using a low-rated load compressor as part of a test apparatus. By reducing the pressure received by the gas turbine engine during its intake and / or exhaust, a verification test can be performed at the maximum combustion temperature and / or maximum cycle pressure ratio of the gas turbine without exceeding the output that the load compressor can absorb. Subsequently, a surge margin test can be performed by introducing compressed air from the load compressor into the combustor of the gas turbine engine.

Description

Technology Field

[0001] The present invention relates to a system and method for off-grid testing of a gas turbine engine, and more specifically, to the use of a system having a low-rated load compressor for performing surge margin testing of a gas turbine having a full-load design output greater than the maximum load of the load compressor. Background Technology

[0002] Gas turbine engines used for power generation (e.g., large onshore gas turbine engines) typically require an extensive set of verification tests to verify that they perform over a wide variety of operating conditions. Historically, most verification tests were performed after the gas turbine engine was designed, manufactured, and installed at a facility, and then connected to the electric grid (i.e., on-grid testing). This requires significant capital expenditure before the gas turbine engine is proven to be effective. The problem to be solved

[0003] The system and method described herein generally provide off-grid verification testing for a gas turbine engine. means of solving the problem

[0004] The system includes a load compressor, an exhaust hood, an intake throttle, and a plurality of fluid conduits. The load compressor is configured to be connected to the rotating shaft of the gas turbine engine and may have a maximum load rating that is smaller than the output generated by the gas turbine engine operating at full speed full load under target operating conditions. The intake throttle is fluidly connected to the intake of the gas turbine engine and configured to control the airflow to the gas turbine engine, thereby controlling the air pressure at the intake of the gas turbine engine. The exhaust hood may be connected to the exhaust port of the gas turbine engine to discharge exhaust gases from the gas turbine engine. A first fluid conduit redirects a first portion of compressed (high-pressure) air or other high-pressure fluid generated by the load compressor to the exhaust hood, causing a pressure drop in the exhaust fluid at the downstream end of the gas turbine engine. In addition, the components of this system simulate operating conditions at higher altitudes, which allows the gas turbine engine to operate over a wider range of operating parameters without exceeding the maximum output that the load compressor can absorb. To perform surge testing, the system also includes a second fluid conduit that returns a second portion of compressed air to the combustor of the gas turbine engine being tested. This enables testing of the engine within the calculated surge margin of the gas turbine engine without exceeding the output or temperature limits of the gas turbine engine. Brief explanation of the drawing

[0005] The above and other aspects will become clearer from the description of the following exemplary embodiments with reference to the attached drawings, in which: FIG. 1 is a perspective view illustrating the interior of a gas turbine engine according to the features described in this specification; FIG. 2 is a compressor map of a gas turbine engine showing the engine's base load operating line, corresponding surge line, and surge margin; FIG. 3 is a schematic diagram of a system according to the features described in this specification; Figure 4 is a comparison compressor map with the gas turbine engine of Figure 2 when the system is used to simulate high-altitude operating conditions; FIG. 5 is a flowchart of a method for a verification test according to the features described in this specification; and FIG. 6 is a flowchart of a surge test method according to the features described in this specification. Examples will be described in detail below with reference to the accompanying drawings. It should be noted that similar numbers refer to similar parts across the various drawings and exemplary embodiments. In some embodiments, detailed descriptions of functions and configurations well known in the art may be omitted so as not to obscure the understanding of the invention by those skilled in the art. For the same reason, some components in the accompanying drawings may be exaggerated, omitted, or schematically depicted. Specific details for implementing the invention

[0006] Before being put into service, gas turbine engines undergo extensive verification testing to provide a reasonable assurance that they will meet expected operating requirements. This verification testing covers various operating modes (e.g., start mode, full speed no load, full speed full load, etc.) and may include the operation of the gas turbine engine under partial load and a wider range of operating conditions than are expected to be encountered in actual operation. Historically, most verification testing is conducted after the gas turbine engine has been designed, manufactured, and installed at the facility. Once installed, the gas turbine engine will be connected to the electric grid (i.e., on-grid testing). This requires a significant capital expenditure before the gas turbine engine is proven to be valid. If the gas turbine engine fails the verification testing, it will be removed and modified before being reinstalled and retested. This requires a significant additional capital expenditure.

[0007] Off-grid testing requires a method to absorb the output generated by a gas turbine engine. One method would be to manufacture a load unit (e.g., dedicated compressor equipment) capable of operating at a level exceeding the maximum output generated by the gas turbine engine. For example, a gas turbine engine with a net output of 100 MW could be coupled to a compressor rated to operate at over 100 MW. However, load compressors of this size are not readily available and are very expensive to manufacture, especially if intended for use solely for verification testing purposes.

[0008] New gas turbine engines are often manufactured to increase the power generation capacity of a power plant. For example, a new gas turbine engine capable of generating more output may be manufactured to replace an older gas turbine engine with a lower rated output. Instead of manufacturing dedicated compressor equipment for the verification testing of the new gas turbine engine, the compressor from the older gas turbine engine may be used instead. Not only is it readily available, but these pre-existing load compressors will also possess well-known performance characteristics. Furthermore, as previously mentioned, compressors from large gas turbine engines may require significant power to operate, making them suitable candidates for this task. However, since the gas turbine engine being tested is typically designed to generate more output than is required to operate the existing load compressor at its maximum load, it cannot be tested using conventional verification test setups across the full operating range of the gas turbine engine, particularly within the upper limit of its design output.

[0009] Additionally, it may be advisable to perform surge testing on the gas turbine engine to be verified as part of the verification test plan. Compressor surges (also known as pressure surges) occur in the engine when compression in the compressor section breaks down, which can cause a reversal of flow so that previously compressed air is expelled out of the engine intake. Compressor surges can occur when the operating cycle pressure ratio exceeds a value above the maximum design cycle pressure ratio. However, performing surge testing can be difficult because this situation occurs outside the normal operating parameters of the gas turbine engine. For example, increasing the cycle pressure ratio of a gas turbine engine can increase the generated output and / or the turbine inlet temperature (TIT) of the gas turbine engine. An increase in TIT can consequently exceed the maximum temperature limits of certain components, while the increased output can exceed the load compressor limits of the verification test setup.

[0010] The present invention solves the problem by simulating the operating conditions of a gas turbine engine that limit the amount of output capable of generating the above problem, while enabling the maximization of other key operating parameters, and by modifying the conditions within the gas turbine engine to induce a compressor surge without exceeding the output limit of an undersized load compressor or the allowable internal temperature of the gas turbine engine.

[0011] First, by redirecting a first portion of the air compressed by the load compressor to a downstream flow within the exhaust hood attached to the gas turbine engine and thereby limiting the air entering the gas turbine engine intake, the system can simulate higher altitude conditions that allow the gas turbine engine to operate at higher compression ratios and temperature conditions without exceeding the output limit of the load compressor.

[0012] Next, by returning a second portion of compressed air from the load compressor to the combustor of the gas turbine engine, the system can be used to perform surge testing of the gas turbine engine. Since the simulation of higher altitude conditions reduces the air pressure downstream of the gas turbine engine, the compressed air from the load compressor has a higher pressure than the air pressure downstream of the gas turbine compressor. When this compressed air is injected into the combustion chamber of the gas turbine engine operating in steady state along its operating line, not only is the mass of air entering the compressor increased, but the back pressure exerted by the gas turbine engine compressor is also increased. Since the output generated by the gas turbine engine is a function of the air mass flow rate, increasing the airflow to the combustion chamber results in a controlled increase in the output generated by the gas turbine engine, allowing the engine to operate within the surge line (i.e., above the operating line but below the surge line) and verifying the engine's ability to operate within the surge margin.

[0013] Surge testing using this method not only prevents the generation of excessive output that could cause the gas turbine engine to overwhelm the low-rated load compressor, but also prevents any other operating parameters (such as the maximum inlet temperature of the gas entering the turbine) from exceeding any limits, such as the melting temperature of certain components or other thermal limits.

[0014] First, referring to FIG. 1, a perspective view of the interior of a gas turbine engine (100) according to one exemplary embodiment is shown. The thermodynamic cycle of the gas turbine engine (100) according to the illustrated embodiment may follow a Brayton cycle. A Brayton cycle may consist of four stages, including isentropic compression (adiabatic compression), constant pressure heating, isentropic expansion (adiabatic expansion), and constant pressure heat dissipation. In a Brayton cycle, thermal energy may be released through the combustion of fuel in a constant pressure environment after the atmosphere is drawn in through an inlet or intake and compressed to high pressure, and then the high-temperature combustion gas is expanded and converted into kinetic energy, and then the exhaust gas containing residual energy is discharged into the atmosphere. That is, a Brayton cycle may consist of four stages: compression, heating, expansion, and exhaust.

[0015] A gas turbine engine (100) using the above Brayton cycle may include a compressor (110), a combustor (120), and a turbine (130) as shown in FIG. 1. The compressor (110) of the gas turbine engine (100) draws in air from the outside and compresses this air. The compressor (110) can supply compressed air to the combustor (120) with compressor blades (113) and can supply cooling air to a high-temperature area for cooling (e.g., components of the turbine (130)). In this case, the air drawn into the compressor (110) undergoes adiabatic compression within it, so the pressure and temperature of the air passing through the compressor (110) rise.

[0016] The compressor (110) can be designed as a centrifugal compressor or an axial compressor. Generally, centrifugal compressors are applied to small gas turbine engines, whereas large gas turbine engines (100) as shown in FIG. 1 require compressing a large amount of air, so multi-stage axial compressors are applied. In a multi-stage axial compressor, the compressor blades (113) of the compressor (110), which rotate along with the rotation of the rotor disks, compress the air introduced therein while transferring the compressed air to the compressor vanes (114) at the rear. The air is gradually compressed to a high pressure as it passes through the compressor blades (113) formed in a multi-stage manner.

[0017] The compressor (110) can be driven by power output from the turbine (130). To this end, as shown in FIG. 1, the rotating shaft (111) of the compressor (110) can be coupled to the rotating shaft (134) of the turbine (130). In a large gas turbine engine (100), nearly half of the output generated from the turbine (130) may be required to drive the compressor (110).

[0018] The combustor (120) can produce high-energy combustion gas by mixing compressed air supplied from the outlet of the compressor (110) with fuel and performing constant-pressure combustion. That is, the combustor (120) can produce high-energy combustion gas by mixing compressed air supplied from the outlet of the compressor (110) with fuel and performing constant-pressure combustion. The combustor (120) includes a plurality of burners located downstream of the compressor (110) and arranged annularly around the central axis of the gas turbine engine (100).

[0019] The turbine (130) comprises a plurality of rotor disks (131) mounted on the rotation axis (134) of the turbine (130), a plurality of turbine blades (132) radially arranged on each of the rotor disks (131), and a plurality of turbine vanes (not shown) located axially upstream of each stage of the turbine blades (132). Each of the rotor disks (131) has a roughly disk shape and has a plurality of grooves formed on its outer periphery. These grooves are each formed with a curved surface so that the turbine blades are inserted into the grooves, and the turbine vanes are mounted within the turbine casing (133). The turbine vanes are fixed so as not to rotate and serve to guide the direction of the flow of combustion gas to the next axially downstream stage of the turbine blades (132). The turbine blades (132) generate rotational force as they rotate due to the combustion gas. The combustion gas exits the gas turbine engine (100) through the exhaust port (140).

[0020] Simply put, the output (PW) generated in the turbine section of a gas turbine engine is the mass flow rate (W) (mass of air + fuel) multiplied by the specific heat of the air / fuel mixture (C p It is the product of ) multiplied by the temperature change (ΔT) across the turbine stage. That is, PW = W x C p x ΔT. The net output of a gas turbine is the output generated in the turbine section minus the work required to operate the rest of the engine (e.g., work required to operate the compressor) and losses due to frictional heat loss, etc.

[0021] By simulating high-altitude conditions, the system reduces the amount of air entering the engine, thereby reducing the total output generated under a set of operating conditions. However, by selectively injecting compressed air into the combustor of a gas turbine engine, the output generated under a set of operating conditions (e.g., specific PR and Nc) can be controlledly increased, causing the system to operate above the operating line and within the surge margin.

[0022] Figure 2 is an exemplary diagram of a compressor map of a gas turbine engine. A typical steady-state operating envelope (T1) of a gas turbine engine is bounded along the upper line of the base load operating line (BLOL). During the transient process, the operating point may temporarily be above the BLOL but below the surge line (SL). The surge line (SL) identifies the point where compressor flow instability increases, where there is a risk of compressor surge. The surge margin (SM) is a quantitative value of the gap from the BLOL to the SL, expressed in terms of compression ratio and flow rate.

[0023] Compressor surges can cause serious damage to the gas turbine engine in addition to interfering with power generation. Accordingly, a verification test system is desirable to confirm whether engine operation within a surge margin across a range of operating conditions will not induce surges. However, this requires a method to reliably increase the amount of output generated under a given set of operating conditions.

[0024] Now, referring to FIG. 3, a verification test system (10) is illustrated. The system (10) may include a gas turbine engine (100) to be tested. The gas turbine engine (100) is coupled to a load compressor (200) and connected to an exhaust hood (400). The gas turbine engine (100) includes an intake (150) through which air is introduced into the compressor (110) of the gas turbine engine (100). The intake (150) is connected and fluidly connected to an air source (600) through an intake duct (610), which in some embodiments may be a filter house. In other embodiments, the air source may be ambient air. An intake throttle (612) is located inside the intake duct (610) upstream of the intake (150) and is configured to regulate the airflow and air pressure as air flows from the air source (600) to the intake (150). In one example, the intake throttle (612) may be the inlet guide vanes (IGVs) of the gas turbine engine itself; in other examples, the intake throttle (612) may be a separate part of the system (10).

[0025] The exhaust hood (400) has a first end (410) that is fluidly connected to the exhaust port (140) of the gas turbine engine (100). In some embodiments, the exhaust hood (400) is connected directly to the exhaust port (140) so as to be in contact with or adjacent to the exhaust port (140). The body (420) of the exhaust hood (400) acts as a channel for exhaust gas exiting from the turbine section of the gas turbine engine (100) to the second end (440) of the exhaust hood (400), where the exhaust gas can safely exit into the surrounding atmosphere. The body (420) may be configured to be approximately straight or to have one or more bends (430), particularly depending on the details of the facility where the system (10) is located. For example, the exhaust hood (400) may have a bend (430) that returns a vertical flow exiting the exhaust hood (400) upward from a horizontal flow roughly aligned with the center axis of the gas turbine engine (100).

[0026] Referring still to FIG. 3, the rotational shaft (134) of the gas turbine engine (100) is coupled to the rotational shaft (240) of the load compressor (200). In one embodiment, the load compressor (200) is a compressor from the existing gas turbine engine. In one embodiment, the rotational shaft (134) of the gas turbine engine (100) and the rotational shaft (240) of the load compressor (200) can be directly connected to each other through a coaxial mechanical coupling so that their rotational movements correspond to each other. In one embodiment, this coupling provides the transfer of rotational energy from the rotational shaft (134) to the rotational shaft (240) when the engine is operating in self-sustain mode or higher. In other embodiments, the rotating shaft (134) of the gas turbine engine (100) and the rotating shaft (240) of the load compressor (200) may be coupled to a starter motor through a torque converter, a gearbox, or a combination thereof, so as to be suitable for transferring rotational energy from the starter motor to the rotating shaft (240) and the rotating shaft (240) during starting.

[0027] The load compressor (200) also further includes an inlet (210) that is fluidly connected to an air source. In one embodiment, the air supplied to the load compressor (200) may originate from an air source (600). An inlet throttle (622) may be located upstream of the inlet (210) and configured to regulate the flow of air supplied from the air source (600) to the inlet (210).

[0028] The load compressor (200) also includes an outlet (220) through which compressed air exits the load compressor (200) when the load compressor is operating. The outlet (220) is fluidly connected to a conduit (300). The conduit (300) includes a first end (302) that can be connected to the outlet (220), and a branch point (310) downstream of the first end, where the conduit (300) is divided into a return conduit (314) and an outlet path (316). A return throttle (312) is located at or adjacent to the branch point (310), which can be used to control the amount of compressed air entering the return conduit (314) and outlet path (316) of the conduit (300). The outlet path (316) can communicate with an outlet that discharges the compressed air into the atmosphere.

[0029] The return conduit (314) subsequently branches into an exhaust hood path (315) and a combustor path (317). The exhaust hood path (315) leads to an entry point (415) of the body (420) of the exhaust hood (400), and enters the exhaust hood cavity (450) of the exhaust hood (400) at a point downstream of the gas turbine engine exhaust port (140). Accordingly, a second conduit is established connecting the exhaust port (outlet 220) of the load compressor to the exhaust hood (400). The exhaust hood path throttle (321) is configured to control the amount of air entering the exhaust hood path (315) from the return conduit (314). The second end (320) of the exhaust hood path (315) is located at a point downstream of the gas turbine engine exhaust port (140) within the exhaust hood (400). The second end (320) of the exhaust hood path (315) may include a nozzle (322), which may be configured to affect the characteristics of the flow of compressed air exiting the second end (320). In one embodiment, the nozzle (322) may be configured to increase the velocity of the compressed air (230) exiting the nozzle (322). In one embodiment, the nozzle (322) is oriented in such a way that the compressed air exiting it is roughly aligned with the downstream direction of the exhaust gas flowing through the body (420) of the exhaust hood (400) at that location within the exhaust hood cavity (450). For example, the nozzle (322) may be oriented in a direction that discharges compressed air perpendicularly from the load compressor (200) to the exhaust hood (400) stack.

[0030] The combustor path (317) is connected to the combustor (120) of the gas turbine engine (100) being tested. Accordingly, a first conduit is established connecting the exhaust port (outlet 220) of the load compressor to the combustor (12) of the gas turbine engine (100). The combustor path throttle (319) is configured to control the flow of air entering the combustor path (317) from the return path (314). The combustor path (317) thus provides a conduit through which a portion of the compressed air is guided from the load compressor (200) to the combustor (120) of the gas turbine engine (100).

[0031] The system (10) may also include an electric starter motor (500) having a starter shaft (510) that is mechanically coupled to the load compressor shaft (240) or the shaft (134) via a releaseable coupling (520). In one embodiment, a torque converter, a gearbox, or a combination thereof may also be used to couple the starter shaft (510) to the load compressor shaft (240) in any manner suitable for the purpose and known in the art. The electric starter motor (500) may be used to start the rotation of the shafts (and connected components) of the gas turbine engine and the load compressor until the gas turbine engine reaches a rotational speed high enough to sustain the operation of the gas turbine engine and the load compressor independently.

[0032] In embodiments where the load compressor (200) is derived from a conventional engine, the performance characteristics of the load compressor (200) will be well known. Alternatively, the load compressor (200) may be a compressor equipment or component specifically developed for use in the system (10). In one embodiment, the full speed-full load (FSFL) based load design output of the gas turbine engine may be greater than the maximum drive output of the load compressor. In one embodiment, the maximum load of the load compressor (200) is 50% to 85% of the FSFL design output of the gas turbine engine (100). In another embodiment, the maximum load of the load compressor (200) is 70% to 75% of the FSFL design output of the gas turbine engine (100).

[0033] The system (10) further includes a control system (700) that is communicably connected to other system components and capable of controlling them, including, without limitation, the control of a gas turbine engine (100) (e.g., control of a fuel injector), a load compressor (200), throttles (612, 622, 312), and an electric starter motor (500). In one embodiment, the control system (700) may include a processor and a memory, wherein the processor may execute commands stored in the memory. In one embodiment, the control system (700) may also be communicably connected to one or more sensors (not shown) on or adjacent to the system (10), the sensors measure other operating parameters of one or more system components, air flow, and / or local environmental factors such as pressure, temperature, air velocity, flow rate, rotational speed, etc., and then provide input to the control system (700).

[0034] In operation, compressed air communicating with the exhaust hood (400) through the intake throttle (612) and the exhaust hood path (315) is used to adjust the volume and pressure of the airflow entering and exiting the gas turbine engine (100). Adjusting the volume and / or pressure in this way simulates the operating conditions that the gas turbine engine (100) may experience at higher altitudes. For example, by using the bypass throttle (312) to introduce compressed air (230) from the load compressor outlet (220) into the conduit (300) and then passing it through the nozzle (322), a pressure drop is induced in the exhaust hood (400), which consequently lowers the pressure in the gas turbine engine exhaust port (140). Injecting the compressed gas from the nozzle (322) at a higher pressure has the effect of lowering the pressure of the exhaust gas exiting the gas turbine engine (100) to below ambient pressure.

[0035] At the same time, by using the intake throttle (612) to reduce the airflow from the air source (600) to the intake duct (610), higher altitude conditions are simulated, resulting in a reduction in the mass of airflow entering the gas turbine engine and a decrease in pressure at the air intake (150). This reduced airflow enables a higher cycle pressure ratio to be tested in the compressor section of the gas turbine engine (100) while generating lower output. By adjusting these parameters, the gas turbine engine (100) can be tested across the main part of its operating envelope without exceeding the maximum load of the load compressor (200).

[0036] Additionally, by using the return conduit (314) and the combustor path (317) to selectively introduce compressed air from the load compressor (200) to the combustor (120) of the gas turbine engine (100), it serves to increase the mass flow of air and fuel into the combustor, which has the effect of increasing the output generated by the gas turbine engine (100) in that specific operating state and provides the capability for a verification test process to test the parameters of the surge line and the accompanying surge margin.

[0037] In the example shown in FIG. 3, the combustor path (317) and the exhaust hood path (315) branch off from the same point of the return conduit (314). Those skilled in the art will readily understand that other configurations of the combustor path (317), the exhaust hood path (315), the return conduit (314), and their respective throttles are also covered within the scope of the invention. For example, in some embodiments, each path (315, 317) may branch off separately from the conduit (300).

[0038] FIG. 4 illustrates a compressor map showing the difference when a gas turbine engine having the compressor map of FIG. 2 operates under simulated high-altitude conditions enabled by the system (10). A comparison between the same velocity line (Nc=1) of the compressor map under sea level and high-altitude conditions shows that the same operating parameters such as PR and Nc have significantly reduced flow rates (W corr It is shown that this results in a reduced output. Accordingly, the surge test by introducing compressed air into the combustor of the gas turbine engine (100) is maintained within the maximum load that can be absorbed by the system's load compressor (200) while generating additional output.

[0039] Referring to FIG. 5, a verification test method (1000) is described here. In step 1010, a starter motor (e.g., an electric starter motor (500)) is used to drive the shafts of a gas turbine engine (e.g., a gas turbine engine (100)) and a load compressor (e.g., a load compressor (200)) to a speed sufficient for the airflow to enable ignition of the combustor. In step 1020, combustion is initiated and the speed of the shaft of the gas turbine engine is increased until a self-sustaining operating speed is reached, after which the starter motor is disengaged from the shaft. Additionally, the gas turbine engine is ramped up to the FSNL operating condition. In step 1030, a control system (e.g., a control system (700)) is subsequently used to adjust one or more operating settings of the system and / or the gas turbine engine so that a first steady-state operation of the gas turbine engine is achieved (e.g., the gas turbine engine is ramped up from the FSNL to a partial load point for testing). The operating settings that can be adjusted include the gas turbine engine air intake (e.g., an intake throttle (612)). It may include settings of system components such as a throttle, a load compressor inlet throttle (e.g., an inlet throttle (622)), a return throttle (e.g., a return throttle (312)), and a nozzle (e.g., a nozzle (322)), and / or settings of a gas turbine engine such as the angle of the gas turbine inlet vane or the amount of fuel injected into the gas turbine engine combustor, among other operating settings. In step 1040, data regarding one or more parameters or performance characteristics of the gas turbine engine is then collected. In step 1050, the operating settings of the system or the gas turbine engine are then adjusted to transition the gas turbine engine to a second steady-state operation.Next, the data collection step 1040 is repeated, and steps 1040 and 1050 are repeated for additional points of steady-state operation for the gas turbine engine until the desired amount of test is completed.

[0040] In examples, the control system is configured so that the operator can adjust a second set of one or more variables (e.g., the amount of fuel injected) while fixing a first set of one or more variables (e.g., air intake pressure), and the control system can automatically adjust certain controls accordingly to achieve a desired steady state. In examples, the control system can be programmed with a predetermined test cycle so that data can be collected over a desired range of operating conditions.

[0041] A surge test method (1100) is illustrated in FIG. 6. To perform a surge test (after or during a verification test of the normal operating envelope of the gas turbine engine), in step 1110, the gas turbine engine is transitioned to normal state operation at a certain speed (e.g., Nc=1) at a point along its operating line (e.g., operating point 1 (OP1) in FIG. 4). Subsequently, in step 1120, the return throttle (e.g., return throttle (312), etc.) and the combustion throttle (e.g., combustion throttle (319), etc.) are adjusted so that compressed air from a load compressor (e.g., load compressor (200), etc.) is introduced into the gas turbine engine combustor (e.g., combustor (120), etc.) through the system's return conduit (e.g., return conduit (314), etc.) and the combustor path (e.g., combustor path (317), etc.) until a surge is possible, imminent, and / or occurs and a surge line is established for a specific point above the operating point (OP1). In step 1130, the compressed air to the combustor is cut off so that the gas turbine engine can reset to normal state operation. In step 1140, the system may then be adjusted until the gas turbine engine establishes a second normal state at a second speed line along the operating line. Then, step 1120 is repeated so that another compressor surge, or near surge, is induced Compressed air is reintroduced into the combustor of the gas turbine engine until the second point of the surge line above this second point of the operating line can be identified. In this way, steps 1120 through 1140 are repeated until the surge line (and its corresponding surge margin) is set along the desired range of operating parameters.

[0042] As mentioned above, compressor surges can be tested until an actual surge occurs, or tested up to a predetermined point above the operating line to prevent potential damage to the gas turbine engine, verifying that no surge occurs below that point. In this way, the surge margin can be verified without exposing the gas turbine engine to repeated compressor surges.

[0043] In one example, the first portion of compressed air directed toward the exhaust hood is less than 50% of the total airflow through the load compressor, and the pressure drop at the outlet is about 30%. In this example, the second portion of the airflow, corresponding to about 20% of the total airflow through the load compressor or within the range of 15% to 25% of the total airflow, is sufficient to increase the cycle pressure ratio of the gas turbine engine to perform surge checks.

[0044] The terms used in this specification are for the purpose of describing specific embodiments only and are not intended to limit the scope of the invention. The singular forms "a," "an," and "the" used in this specification are intended to include the plural forms as well, unless the context clearly indicates otherwise. In the invention, terms such as "comprises," "includes," or "have / has" specify the presence of such features, figures, steps, operations, components, parts, and / or combinations thereof, but should be interpreted as not excluding the presence or addition of one or more other features, figures, steps, operations, components, parts, and / or combinations thereof.

[0045] (Although) one or more examples have been described with reference to the attached drawings, it will be obvious to those skilled in the art that variations and modifications can be made by adding, changing, or removing components without departing from the concept and scope of the invention as defined in the attached claims, and such variations and modifications are encompassed within the concept and scope of the invention as defined in the attached claims.

Claims

Claim 1 A system for off-grid testing of a gas turbine engine using a low-rated load compressor, wherein the system comprises: a gas turbine engine including an intake, a compressor, a combustor, a turbine, and a rotating shaft; a load compressor mechanically coupled to the rotating shaft of the gas turbine engine and having an exhaust port; a first conduit fluidly connecting the exhaust port of the load compressor to the combustor of the gas turbine engine; an exhaust hood fluidly connected to the gas turbine engine downstream of the turbine so as to discharge exhaust gas discharged from the turbine to the exhaust hood; and a second conduit fluidly connecting the exhaust port of the load compressor to the exhaust hood. Claim 2 A system according to claim 1, further comprising a first throttle configured to be fluidly connected to the first conduit and to limit a first portion of the flow of compressed air from the load compressor to the first conduit. Claim 3 In paragraph 2, a system in which the gas turbine engine has a full-speed full-load (FSFL) design output and the load compressor has a maximum load less than the FSFL design output. Claim 4 A system further comprising a second throttle configured to be fluidly connected to the second conduit and to restrict a second portion of the flow of compressed air from the load compressor to the second conduit in paragraph 3. Claim 5 A system according to claim 4 further comprising a third throttle configured to be fluidly connected to the intake of the gas turbine engine and to restrict the flow of intake air into the intake. Claim 6 A system according to claim 5, further comprising a control system configured to operate the gas turbine engine, the load compressor, the first throttle, the second throttle, and the third throttle. Claim 7 In paragraph 6, the control system comprises a processor and a memory that stores a set of operation instructions. Claim 8 A system according to claim 7, further comprising at least one sensor configured to communicate with the processor and measure at least one parameter of the gas turbine engine. Claim 9 In claim 8, the system wherein the processor commands the adjustment of at least one operating parameter of the system according to a first set of operating commands based on input received from at least one sensor. Claim 10 A system according to claim 1, further comprising a starter motor mechanically coupled to the rotating shaft of the gas turbine engine. Claim 11 A method for conducting off-grid testing of a gas turbine engine using a low-rated load compressor, the method comprising: a step of mechanically coupling a first rotating shaft of the gas turbine engine to a second rotating shaft of the low-rated load compressor, wherein the gas turbine engine has a full-force full load (FSFL) design output and the low-rated load compressor has a maximum load smaller than the FSFL design output of the gas turbine engine; a step of simulating the operating conditions of the gas turbine engine using a first portion of the total flow of compressed air generated by the load compressor; a step of inducing the gas turbine engine to approach a compressor surge operating state beyond the design operating state; and a step of generating test data based on the gas turbine engine operating beyond the design operating state. Claim 12 A method according to claim 11, wherein the step of simulating the operating conditions comprises: a step of setting a first pressure lower than ambient atmospheric pressure within the exhaust hood fluidly connected downstream of the gas turbine engine by introducing a first portion of the total flow of the compressed air into the exhaust hood; and a step of setting an air intake pressure corresponding to the first pressure in the air intake of the gas turbine engine. Claim 13 A method according to claim 12, wherein the induction step comprises introducing a second portion of the total flow of compressed air generated by the low-rated load compressor into the combustor of the gas turbine engine. Claim 14 In paragraph 13, the method wherein the second portion is 15% to 25% of the total flow of compressed air generated by the low-rated load compressor. Claim 15 In paragraph 14, the method in which the first portion is less than 50% of the total flow of compressed air generated by the low-rated load compressor. Claim 16 A method according to claim 13 further comprising the steps of: ceasing to introduce the second portion of the total flow of compressed air generated by the low-rated load compressor into the combustor; resetting the design operating state; setting the second design operating state; and inducing the gas turbine engine to operate in a manner close to the second compressor surge operating state beyond the second design operating state. Claim 17 A method according to claim 11 in which the maximum load of the low-rated load compressor is 70% to 75% of the FSFL design output of the gas turbine engine. Claim 18 A method according to claim 11 in which the maximum load of the low-rated load compressor is 50% to 85% of the FSFL design output of the gas turbine engine. Claim 19 A system for off-grid testing a gas turbine engine having a full-load ("FSFL") design output, wherein the gas turbine engine has a compressor having an air intake, a combustor, a turbine, and a rotating shaft connecting the turbine to the compressor, the system comprises: a load compressor having an exhaust port configured to be mechanically connected to the rotating shaft of the gas turbine engine having a maximum load smaller than the FSFL design output of the gas turbine engine; an intake throttle configured to be fluidly connected to the air intake of the gas turbine engine and to reduce the pressure of air entering the air intake; an exhaust hood configured to be connected to the turbine downstream of the gas turbine engine so that exhaust gas communicates from the turbine to the exhaust hood; an injection nozzle fluidly connected to the exhaust hood; a first bypass conduit fluidly connecting the exhaust port of the load compressor to the combustor of the gas turbine engine; and a second bypass conduit fluidly connecting the exhaust port of the load compressor to the injection nozzle to communicate high-pressure fluid to the exhaust hood. Claim 20 A system according to claim 19 further comprising a first throttle configured to control the flow of fluid through the first bypass conduit and a second throttle configured to control the flow of fluid through the second bypass conduit.

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