CN117969329A - Method for testing accelerated thermal shock of high-pressure turbine guide vane of marine gas turbine - Google Patents
Method for testing accelerated thermal shock of high-pressure turbine guide vane of marine gas turbine Download PDFInfo
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Abstract
The invention discloses a high-pressure turbine guide vane acceleration thermal shock test method of a marine gas turbine, which comprises the following steps of: step 1, determining the aero-thermal parameters of a test blade; step 2, determining the number of single-group test blades; step 3, determining the number of test groups; step 4, determining a temperature load spectrum curve of the high-pressure turbine guide vane; step 5, debugging the temperature of the blade; step 6, performing a formal acceleration thermal shock test; step 7, evaluating the state of the high-pressure turbine guide vane test blade; step 8, determining the thermal fatigue cycle resistance of the high-pressure turbine guide vane; step 9, performing simulation calculation on the thermal fatigue resistance circulation capacity of the high-pressure turbine guide vane; and 10, analyzing the validity of the test result. According to the working characteristics of frequent start and stop of the high-pressure turbine of the marine gas turbine, the test process is completely organized, and the thermal fatigue resistance assessment of the high-pressure turbine guide vane can be realized with lower technical risk and investment and near real working environment conditions under the state of parts.
Description
Technical Field
The invention belongs to the field of marine turbines, and particularly relates to a high-pressure turbine guide vane acceleration thermal shock test method of a marine gas turbine.
Background
The gas turbine has the advantages of high power density, high starting speed, flexible fuel and the like, and is widely applied to the fields of industrial and offshore platform power generation, natural gas transportation, petrochemical industry, metallurgy and the like, and is also widely used as a main power device of a ship.
Modern high performance gas turbines are continually increasing in gas initial temperature (high pressure turbine inlet temperature) for higher cycle efficiency, higher power. With the increasing inlet temperature of the high-pressure turbine, the operation temperature of the high-pressure turbine is far higher than the melting point temperature of the blade materials, such as the inlet gas temperature of the turbine of the most advanced gas turbine which is put into operation at present reaches 1600 ℃, and the inlet gas temperature of the turbine of the advanced aero-engine is more higher than 1800 ℃. There are three main measures to ensure that a gas turbine blade can safely and reliably operate for a long period of time in such a high temperature environment: firstly, the heat-resistant grade of the turbine blade material is continuously improved, secondly, an advanced cooling technology is adopted to reduce the temperature of the blade, and thirdly, the heat-insulating effect of the heat-insulating coating of the turbine blade is continuously improved. In recent years, the increase in turbine inlet temperature has been mainly due to the increase in turbine cooling design level, and secondly due to the development of high-performance heat-resistant alloys and coating materials and the progress of the production and manufacturing process level. Obviously, turbine blade cooling plays a vital role in increasing turbine inlet temperature and improving gas turbine performance. But this also results in higher and higher turbine blade thermal loads, which presents a greater challenge for the reliability of the blade.
In addition, in the running and using process of the marine gas turbine, the operations such as starting, accelerating, decelerating and stopping lead the high-pressure turbine blade to bear larger temperature load change, namely thermal shock load, and the frequent and repeated temperature load change is extremely easy to lead the high-pressure turbine blade to be in thermal shock fatigue, so that the service life of the high-pressure turbine blade is greatly reduced, and the reliability and the safety of a unit are influenced. In addition, the special marine salt fog working environment and the influence of sulfur element in fuel oil become non-negligible factors in the design and test of marine gas turbines.
In recent years, although scholars and researchers at home and abroad have conducted a great deal of research on the aspect of efficient cooling design of turbine blades, have conducted a small amount of research on the aspect of thermal shock test of turbine blades of aeroengines, and have conducted a certain knowledge on improving the cooling performance of turbine blades and revealing the cooling flow mechanism inside the blade bodies of turbine blades, the research has not focused on how to conduct a thermal shock test on a marine gas turbine so as to improve the thermal fatigue resistance of high-pressure turbine blades of the marine gas turbine, and has been reported on the aspect of thermal shock test of turbine blades of the marine gas turbine. Therefore, the thermal fatigue test and examination of the high-pressure turbine blade has extremely important practical value and scientific significance, and particularly, how to scientifically, reasonably and pointedly develop the thermal fatigue test and examination of the high-pressure turbine guide vane of the marine gas turbine for starting and stopping the high-pressure turbine for use under all-condition frequent load change becomes a technical problem to be solved in urgent need of the development of the marine gas turbine.
Disclosure of Invention
In order to solve the technical problems, the invention provides a method for accelerating thermal shock test of a high-pressure turbine guide vane of a marine gas turbine, which is used for realizing the assessment of the thermal fatigue resistance of the high-pressure turbine guide vane under the condition of low technical risk and investment and near the real working environment.
The invention aims at realizing the following technical scheme, and discloses a high-pressure turbine guide vane acceleration thermal shock test method of a marine gas turbine, which comprises the following steps of:
Step 1, determining the aero-thermal parameters of a test blade;
step2, determining the number of single-group test blades;
Step 3, determining the number of test groups;
step 4, determining a temperature load spectrum curve of the high-pressure turbine guide vane;
step 5, debugging the temperature of the blade;
step 6, performing a formal acceleration thermal shock test;
step 7, evaluating the state of the high-pressure turbine guide vane test blade;
step 8, determining the thermal fatigue cycle resistance of the high-pressure turbine guide vane;
Step 9, performing simulation calculation on the thermal fatigue resistance circulation capacity of the high-pressure turbine guide vane;
And 10, analyzing the validity of the test result.
Preferably, in step 2, the number of the single group of test blades is determined according to the ratio of the maximum gas flow of the turbine guide blade thermal shock test stand to the gas flow of the single turbine blade cascade channel in the gas thermal parameters of the test blades.
Preferably, in step3, the number of test groups is determined according to the following formula:
Wherein, N s,s is the number of single group of test blades, and N s,all is the total number of high-pressure turbine guide vanes required to carry out thermal shock test.
Preferably, in step 4, the high pressure turbine vane temperature load profile is determined according to the following rules: the quick heating duration is not larger than the time used when the marine gas engine is quickly increased from the empty load to the full load, the highest temperature is equal to the average temperature of the middle section of the high-pressure turbine guide vane under the full load condition, the highest temperature duration is determined according to the time when the blade reaches the stable highest temperature under the full load condition, the quick cooling duration is not larger than the time used when the gas engine is quickly reduced from the full load to the empty load, the lowest temperature is the minimum value of the average temperature of the middle section of the high-pressure turbine guide vane under the empty load condition and the lowest average temperature which can be reached by the section of the high-pressure turbine guide vane under the lowest stable combustion condition of the test bed combustor, and the lowest temperature duration is determined according to the time when the blade reaches the stable lowest temperature under the empty load condition.
Preferably, in step 5, the blade temperature commissioning test comprises the steps of: selecting a certain high-pressure turbine guide vane, arranging thermocouples in the middle section of the vane body of the high-pressure turbine guide vane as debugging vanes, installing the debugging vanes on a test bed, adjusting the oil supply flow of a burner of the test bed, the air supply flow of the burner, the air supply pressure of the cooling air and the air supply flow of the cooling air according to the temperature load spectrum curve of the high-pressure turbine guide vane given in the step 4, ensuring that the highest temperature and the lowest temperature load states in the temperature load spectrum of the high-pressure turbine guide vane can be realized, recording parameters of the oil supply flow of the burner, the air supply pressure of the cooling air and the air supply flow of the cooling air of the test bed under the two load states, and performing test state control according to the parameters during the follow-up formal test.
Preferably, in step 6, the formally accelerated thermal shock test comprises the steps of: photographing and recording the state of the high-pressure turbine guide vane before the test of the test vane; detaching the debugging blade and replacing the debugging blade with a high-pressure turbine guide blade without a thermocouple; and (3) on the basis of completing the blade temperature debugging test in the step (5), developing a high-pressure turbine guide vane acceleration thermal shock test according to the high-pressure turbine guide vane temperature load spectrum curve determined in the step (4) according to parameters of the combustor oil supply flow, the combustor air supply pressure, the cooling air supply pressure and the cooling air supply flow under the two load states of the highest temperature and the lowest temperature determined in the debugging test.
Preferably, in step 7, the high pressure turbine vane test blade state assessment includes the steps of: in the accelerated thermal shock test process, after 500 cycles, checking and analyzing the surface state of the high-pressure turbine guide vane test blade, photographing and recording the state of the high-pressure turbine guide vane test blade after the current cycle times, and if cracks with the length exceeding 1mm appear on the surface, ending the test, wherein the current completed cycle times are the thermal fatigue resistance cycle capacity of the test blade; if no cracks appear on the surface, the test is continued until cracks exceeding 1mm in length appear.
Preferably, in the step 8, the states of the high-pressure turbine guide vane test blades recorded by each inspection, analysis and photographing before the test in the step 6 and in the test in the step 7 are compared and analyzed, and the thermal fatigue resistance circulation capacity of the high-pressure turbine guide vane is comprehensively determined by combining the thermal fatigue resistance circulation capacity of the test blades determined by the inspection, analysis and determination in the test in the step 7.
Preferably, in step 9, the high pressure turbine vane thermal fatigue resistance cycle capability simulation calculation includes the steps of: inputting the temperature load spectrum determined in the step 4 into a simulation calculation model for calculation, and respectively obtaining equivalent strain distribution corresponding to the peak temperature and the valley temperature T test,min, wherein the equivalent strain calculated under the working condition of the peak temperature T test,max of the examination point of the examination section is recorded asEquivalent strain corresponding to valley temperature T test,min of examination section examination point is recorded as/>Its equivalent strain range is/>Bringing it into the following formula to obtain the number of cycles N f as shown in formula (2):
Wherein b is a fatigue strength index, c is a fatigue ductility index, σ f is a fatigue strength coefficient, ε f is a fatigue ductility coefficient, E is an elastic modulus, and N f is a thermal fatigue cycle resistance.
Preferably, in step 10, comparing the thermal fatigue resistance cycling ability N f obtained by calculation in step 9 with N test obtained by test in step 8, if N f/Ntest is greater than or equal to 5, N test is the thermal fatigue resistance cycling ability of the high pressure turbine guide vane; if N f/Ntest is less than 5, the test protocol needs to be readjusted for testing.
Compared with the prior art, the invention has the following advantages:
According to the method for testing the accelerated thermal shock of the high-pressure turbine guide vane of the marine gas turbine, provided by the invention, according to the working characteristics of frequent start and stop of the high-pressure turbine of the marine gas turbine, the test process is completely new, the method is beneficial to standardizing the accelerated thermal shock test process of the high-pressure turbine guide vane of the marine gas turbine, and the high-pressure turbine guide vane of the marine gas turbine meeting the use requirement can be obtained. By adopting the test method provided by the invention, the high-pressure turbine guide vane thermal fatigue resistance can be checked under the condition of parts with lower technical risk and investment and under the condition of approaching to a real working environment, a complex complete machine environment is not needed, and a large number of complete machine up-and-down test benches and high manpower and material resource investment for decomposition and inspection are avoided.
Drawings
FIG. 1 is a flow chart of a method for testing the accelerated thermal shock of the high pressure turbine guide vanes of a marine gas turbine in an embodiment of the invention;
FIG. 2 is a graph of the load spectrum determined in step 4 in an embodiment of the present invention.
Detailed Description
The invention is described in further detail below with reference to the drawings and examples. It is to be understood that the specific embodiments described herein are merely illustrative of the invention and are not limiting thereof. It should be further noted that, for convenience of description, only some, but not all of the structures related to the present invention are shown in the drawings.
As shown in fig. 1, the technical scheme of the invention provides a method for accelerating thermal shock test of a high-pressure turbine guide vane of a marine gas turbine, which comprises the following steps:
Step 1, determining the aero-thermal parameters of the test blade: according to the temperature field results given by the design and calculation of the high-pressure turbine guide vane, the average temperature T s,f,ave of the middle section of the high-pressure turbine guide vane under the condition of full load, the average temperature T s,e,ave of the middle section of the high-pressure turbine guide vane under the condition of empty load, the gas flow G s,g of a single turbine blade cascade channel under the test condition and the total gas temperature are given Total pressure of fuel gas/>Total pressure of cooling air/>Total temperature of cooling air/>The single high pressure turbine vane cooling air flow G s,c.
Step 2, determining the number of single-group test blades: according to the gas flow G s,g of the single turbine blade cascade channel in the gas-heat parameters of the test blades determined in the step 1, combining the maximum gas flow G T,g,max of a turbine guide blade thermal shock test bed, and if G T,g,max/Gs,g is more than or equal to 4, taking 3 from the number N s,s of the single set of test blades; if G T,g,max/Gs,g is more than or equal to 3 and less than 4, the number N s,s of the single group of test blades is 2; if 2.ltoreq.G T,g,max/Gs,g <3, the number of the single set of test leaves N s,s is 1.
Step 3, determining the number of test groups: the number of test groups was determined according to the following formula:
Wherein N s,s is the number of single set of test blades, determined by step 2, N s,all is the total number of high pressure turbine vanes required to develop a thermal shock test.
Step 4, determining a high-pressure turbine guide vane temperature load spectrum curve: according to the average temperature T s,f,ave of the middle section of the high-pressure turbine guide vane under the full load condition and the average temperature T s,e,ave of the middle section of the high-pressure turbine guide vane under the empty load condition of the test blade gas-heat parameter determined in the step 1, the lowest average temperature T s,Tmin,ave of the middle section of the high-pressure turbine guide vane under the state of lowest stable combustion can be achieved by combining with the lowest outlet temperature T comb,min of the test bed burner, and the temperature load spectrum of the high-pressure turbine guide vane is given: the quick heating duration T test,heat is not more than the time T load,0-1 used when the marine gas turbine quickly rises from the air load to the full load, the highest temperature T test,max＝Ts,f,ave, the highest temperature duration T test,max is determined according to the time when the blade reaches the stable highest temperature in the full load state, the highest temperature duration T test,max =2-3 s is general, the quick cooling duration T test,cool is not more than the time T load,1-0 used when the gas turbine quickly falls from the full load to the air load, the lowest temperature T test,min＝min(Ts,e,ave,Ts,Tmin,ave), the lowest temperature duration T test,min is determined according to the time when the blade reaches the stable lowest temperature in the air load state, the lowest temperature duration T test,min =2-3 s is general, and a load spectrum curve is drawn as shown in fig. 2.
Step 5, debugging the temperature of the blade: installing debugging blades with thermocouples arranged on the middle sections of the blade bodies of the high-pressure turbine guide vanes on a test bed, and adjusting the oil supply flow G T,oil, the gas supply flow G T,g and the gas supply pressure of the burner of the test bed according to the temperature load spectrum curve of the high-pressure turbine guide vanes given in the step 4Cooling air supply pressure/>The cooling air supply flow G T,c ensures that two load states (the deviation is not more than 1%) of the highest temperature T test,max and the lowest temperature T test,min in a high-pressure turbine guide vane temperature load spectrum can be realized, and the oil supply flow G T,oil, the gas supply flow G T,g and the gas supply pressure/>, of the test bed burner under the two load states are recordedCooling air supply pressure/>And the cooling air supply flow G T,c is a parameter, and the test state is controlled according to the parameter in the follow-up formal test.
And step 6, performing a formal accelerated thermal shock test: photographing and recording the state of the high-pressure turbine guide vane before the test of the test vane; detaching the debugging blade with the thermocouple installed in the step 5, and replacing the debugging blade with the high-pressure turbine guide blade without the thermocouple; on the basis of completing the blade temperature debugging test in the step 5, according to the maximum temperature T test,max and the minimum temperature T test,min determined by the debugging test, the burner oil supply flow G T,oil, the burner air supply flow G T,g and the burner air supply pressure under two load statesCooling air supply pressure/>And (3) carrying out a high-pressure turbine guide vane acceleration thermal shock test according to the high-pressure turbine guide vane temperature load spectrum curve determined in the step (4) by using the cooling air supply flow G T,c.
Step 7, evaluating the state of the high-pressure turbine guide vane test blade: in the test process of step 6, after 500 cycles, checking and analyzing the surface state of the high-pressure turbine guide vane test blade, photographing and recording the state of the high-pressure turbine guide vane test blade after the current cycle times, and if cracks with the length exceeding 1mm appear on the surface, ending the test, wherein the current completed cycle times are the thermal fatigue resistance cycle capacity of the test blade; if no crack appears on the surface, the test is continued until the crack with the length exceeding 1mm appears, and the cycle number is the thermal fatigue resistance cycle capability of the test blade.
Step 8, determining the thermal fatigue cycle resistance of the high-pressure turbine guide vane: and (3) comparing and analyzing the states of the high-pressure turbine guide vane test blades recorded by each examination, analysis and photographing before the test in the step (6) and in the test process in the step (7), and finally determining the thermal fatigue cycle resistance N test of the high-pressure turbine guide vane by combining the thermal fatigue cycle resistance of the test blades determined by the examination, analysis and determination in the test process in the step (7).
Step 9, simulating and calculating the thermal fatigue resistance circulation capacity of the high-pressure turbine guide vane: inputting the temperature load spectrum determined in the step 4 into a simulation calculation model for calculation, and respectively obtaining equivalent strain distribution corresponding to the peak temperature T test,max and the valley temperature T test,min, wherein the equivalent strain calculated under the working condition of the peak temperature T test,max of the examination point of the examination section is recorded asEquivalent strain corresponding to valley temperature T test,min of examination section examination point is recorded as/>The equivalent strain range isThe cycle times N f are obtained by bringing the heat fatigue resistance performance of the high-pressure turbine guide vane into the following formula, and the heat fatigue resistance performance of the high-pressure turbine guide vane is shown as the formula (2):
Wherein b is a fatigue strength index, c is a fatigue ductility index, σ f is a fatigue strength coefficient, ε f is a fatigue ductility coefficient, E is an elastic modulus, and N f is a thermal fatigue cycle resistance.
Step 10, test result validity analysis: comparing the thermal fatigue resistance cycling ability N f obtained by calculation in the step 9 with N test obtained by the test in the step 8, and if N f/Ntest is more than or equal to 5, obtaining N test as the thermal fatigue resistance cycling ability of the high-pressure turbine guide vane; if N f/Ntest is less than 5, the test protocol needs to be readjusted for testing.
The foregoing is a preferred embodiment of the present invention and it should be noted that modifications and adaptations to those skilled in the art may be made without departing from the principles of the present invention and are intended to be comprehended within the scope of the present invention.
Claims (10)
1. A method for testing the accelerated thermal shock of a high-pressure turbine guide vane of a marine gas turbine is characterized by comprising the following steps of: the method comprises the following steps:
Step 1, determining the aero-thermal parameters of a test blade;
step2, determining the number of single-group test blades;
Step 3, determining the number of test groups;
step 4, determining a temperature load spectrum curve of the high-pressure turbine guide vane;
step 5, debugging the temperature of the blade;
step 6, performing a formal acceleration thermal shock test;
step 7, evaluating the state of the high-pressure turbine guide vane test blade;
step 8, determining the thermal fatigue cycle resistance of the high-pressure turbine guide vane;
Step 9, performing simulation calculation on the thermal fatigue resistance circulation capacity of the high-pressure turbine guide vane;
And 10, analyzing the validity of the test result.
2. The method for accelerating thermal shock test of a high pressure turbine vane of a marine gas turbine as set forth in claim 1, wherein: in the step 2, the number of the single group of test blades is determined according to the ratio of the maximum gas flow of the turbine guide blade thermal shock test bed to the gas flow of the single turbine blade cascade channel in the gas thermal parameters of the test blades.
3. The method for accelerating thermal shock test of a high pressure turbine vane of a marine gas turbine as set forth in claim 1, wherein: in the step 3, the number of test groups is determined according to the following formula:
Wherein, N s,s is the number of single group of test blades, and N s,all is the total number of high-pressure turbine guide vanes required to carry out thermal shock test.
4. The method for accelerating thermal shock test of a high pressure turbine vane of a marine gas turbine as set forth in claim 1, wherein: in the step 4, the temperature load spectrum curve of the high-pressure turbine guide vane is determined according to the following rule: the quick heating duration is not larger than the time used when the marine gas engine is quickly increased from the empty load to the full load, the highest temperature is equal to the average temperature of the middle section of the high-pressure turbine guide vane under the full load condition, the highest temperature duration is determined according to the time when the blade reaches the stable highest temperature under the full load condition, the quick cooling duration is not larger than the time used when the gas engine is quickly reduced from the full load to the empty load, the lowest temperature is the minimum value of the average temperature of the middle section of the high-pressure turbine guide vane under the empty load condition and the lowest average temperature which can be reached by the section of the high-pressure turbine guide vane under the lowest stable combustion condition of the test bed combustor, and the lowest temperature duration is determined according to the time when the blade reaches the stable lowest temperature under the empty load condition.
5. The method for accelerating thermal shock test of high-pressure turbine guide vanes of a marine gas turbine according to claim 4, wherein: in the step 5, the blade temperature debugging test comprises the following steps: selecting a certain high-pressure turbine guide vane, arranging thermocouples in the middle section of the vane body of the high-pressure turbine guide vane as debugging vanes, installing the debugging vanes on a test bed, adjusting the oil supply flow of a burner of the test bed, the air supply flow of the burner, the air supply pressure of cooling air and the air supply flow of the cooling air according to the temperature load spectrum curve of the high-pressure turbine guide vane given in the step 4, ensuring that two load states of the highest temperature and the lowest temperature in the temperature load spectrum of the high-pressure turbine guide vane can be realized, recording parameters of the oil supply flow of the burner, the air supply pressure of the cooling air and the air supply flow of the cooling air of the test bed under the two load states, and performing test state control according to the parameters during the follow-up formal test.
6. The method for accelerating thermal shock test of high-pressure turbine guide vanes of a marine gas turbine according to claim 5, wherein: in the step 6, the formal accelerated thermal shock test comprises the following steps: photographing and recording the state of the high-pressure turbine guide vane before the test of the test vane; detaching the debugging blade and replacing the debugging blade with a high-pressure turbine guide blade without a thermocouple; and on the basis of completing the blade temperature debugging test in the step 5, developing a high-pressure turbine guide vane acceleration thermal shock test according to the high-pressure turbine guide vane temperature load spectrum curve determined in the step 4 according to the parameters of the combustor oil supply flow, the combustor air supply pressure, the cooling air supply pressure and the cooling air supply flow under the two load states of the highest temperature and the lowest temperature determined in the debugging test.
7. The method for accelerating thermal shock test of a high pressure turbine vane of a marine gas turbine as set forth in claim 6, wherein: in the step 7, the high-pressure turbine guide vane test blade state evaluation comprises the following steps: in the accelerated thermal shock test process, after 500 cycles, checking and analyzing the surface state of the high-pressure turbine guide vane test blade, photographing and recording the state of the high-pressure turbine guide vane test blade after the current cycle times, and if cracks with the length exceeding 1mm appear on the surface, ending the test, wherein the current completed cycle times are the thermal fatigue resistance cycle capacity of the test blade; if no cracks appear on the surface, the test is continued until cracks exceeding 1mm in length appear.
8. The method for accelerating thermal shock test of a marine gas turbine high pressure turbine vane as set forth in claim 7, wherein: in the step 8, the states of the high-pressure turbine guide vane reference blades recorded by each inspection, analysis and photographing before the test in the step 6 and in the test in the step 7 are compared and analyzed, and the thermal fatigue resistance cycle capacity of the high-pressure turbine guide vane is comprehensively determined by combining the thermal fatigue resistance cycle capacity of the test blades determined by the inspection, analysis and determination in the test in the step 7.
9. The method for accelerating thermal shock test of a high pressure turbine vane of a marine gas turbine as set forth in claim 8, wherein: in the step 9, the simulation calculation of the thermal fatigue resistance circulation capacity of the high-pressure turbine guide vane comprises the following steps: inputting the temperature load spectrum determined in the step 4 into a simulation calculation model for calculation, and respectively obtaining equivalent strain distribution corresponding to the peak temperature and the valley temperature T test,min, wherein the equivalent strain calculated under the working condition of the peak temperature T test,max of the examination point of the examination section is recorded asEquivalent strain corresponding to valley temperature T test,min of examination section examination point is recorded as/>The equivalent strain range isBringing it into the following formula to obtain the number of cycles N f as shown in formula (2):
Wherein b is a fatigue strength index, c is a fatigue ductility index, σ f is a fatigue strength coefficient, ε f is a fatigue ductility coefficient, E is an elastic modulus, and N f is a thermal fatigue cycle resistance.
10. The method for accelerating thermal shock test of a marine gas turbine high pressure turbine vane as set forth in claim 9, wherein: in the step 10, comparing the thermal fatigue resistance cycling ability N f obtained by calculation in the step 9 with N test obtained by the test in the step 8, if N f/Ntest is more than or equal to 5, N test is the thermal fatigue resistance cycling ability of the high-pressure turbine guide vane; if N f/Ntest is less than 5, the test protocol needs to be readjusted for testing.
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