رقم المقالة : JSME-2310-1273 (R1) زيارة : 167 الصفحة: 5 - 21

نوع المخطوط: ابحاث

Conceptual design of the inlet flow preheat system to the nozzle test equipment of a turbine engine

الموضوعات : فصلنامه شبیه سازی و تحلیل تکنولوژی های نوین در مهندسی مکانیک

Behrooz Shahriari ¹ , Hamid Farrokhfal ² , Mohammad Reza Nazari ³

1 - Faculty of Mechanics, Malek Ashtar University of Technology, Isfahan, Iran
2 - Faculty of Mechanics, Malek Ashtar University of Technology, Iran
3 - Faculty of Mechanics, Malek Ashtar University of Technology, Iran

تاريخ الإرسال : 16 الأحد , ربيع الأول, 1445 تاريخ التأكيد : 10 الأحد , ذو الحجة, 1445 تاريخ الإصدار : 18 الأربعاء , محرم, 1446

الکلمات المفتاحية: Turbofan engine nozzle, Nozzle ground test, Hot flow simulation, Combustion chamber design, Afterburner design,

ملخص المقالة :

To ensure the correct design and proper operation of turbine engine components, various types of ground tests must be performed, which requires the simulation of the input flow to these components. To test hot components such as nozzles, it is necessary to create a hot flow, which is done by the preheat system of the inlet flow. For the ground test of the nozzle of a turbofan engine that has an afterburner in addition to the combustion chamber, the preheat system must be able to supply air in two modes, dry mode and reheat mode. In this paper a combustion chamber with an afterburner is used to provide hot air to the nozzle. In the first step, using experimental and analytical relations, the combustion chamber is designed. The presented algorithm has the ability to calculate the diameter and reference surface of the combustion chamber, components of the combustion chamber and thermodynamic parameters. Then the obtained results are compared with the data of a similar annular combustion chamber. The comparison indicates the acceptable convergence of the design results with the experimental results. Finally, the output flow of the combustion chamber is considered as the input of the afterburner, and the temperature of the combustion flow is increased by the afterburner to the desired temperature. In addition to the ability to design a V-gutter flame keeper, the afterburner design algorithm also calculates the thermodynamic characteristics of the afterburner output stream, considering the requirements of flame stability.

المصادر:

[1] W. Dodds and D. Bahr. (1990). Combustion system design. Design of modern gas turbine combustors, pp.343-476.
[2] J. D. Mattingly. (2002). Aircraft engine design: AIAA.
[3] A. H. Lefebvre. (2010). Gas turbine combustion. CRC Press.
[4] M. R. J. Charest. (2006). Design methodology for a lean premixed prevaporized can combustor. MS Thesis, Library and Archives Canada,Carleton University, Department of Mechanical and Aerospace Engineering.
[5] P. J. Stuttaford and P. A. Rubini. (1996). Preliminary gas turbine combustor design using a network approach. In ASME International Gas Turbine and Aeroengine Congress and, Exhibition.
[6] P. J. Stuttaford and P. A. Rubini. (1996) Preliminary gas turbine combustor design using a network approach. in ASME International Gas Turbine and Aeroengine Congress and, Exhibition.
[7] M. J. Wankhede. (2012). Multi-fidelity strategies for lean burn combustor design. PhD Thesis, University of Southampton.
[8] Jai-Houng Leu. (2010). Design and simulation validation of LHV fuel combustor. Journal of Information and Optimization Sciences, 31:6,1321-1336.
[9] Yize Liu, Xiaoxiao Sun, Vishal Sethi, Yi-Guang Li, Devaiah Nalianda, David Abbott, Pierre Gauthier,Bairong Xiao and Lu Wang. (2021). Development and application of a preliminary design methodology for modern low emissions aero combustors. Journal of power and energy. Vol. 235(4) 783-806.
[10] J. W. Sawyer. (1966). Gas Turbine Engineering Handbook: Editor. John W. Sawyer, vol. , Gas Turbine Publications.
[11] P. P. Walsh and P. Fletcher. (2004). Gas turbine performance. John Wiley & Sons.
[12 ] R. Rezvani. (2010). A Conceptual Methodology for the Prediction of Engine Emissions. PhD Thesis, Georgia Tech. University.
[13] H. Knight and R. Walker. (1953). The component pressure losses in combustion chambers. Gt. Brit. National Gas Turbine Establishment, Farnborough, Hants, England.
[14] M. R. J. Charest. (2006). Design methodology for a lean premixed prevaporized can combustor. MSThesis, Library and Archives Canada Bibliothèque et Archives Canada, Carleton University,
Department of Mechanical and Aerospace Engineering.
[15] A. Costa Conrado, P. T. Lacava, A. C. P. Filaho, M. D. S. Sanches. (2004). Basic design principles for gas turbine combustor. Proceedings of the 10o Brazilian Congress of Thermal Sciences and Engineering.
[16] J. J. Gouws. (2008). Combining a one-dimensional empirical and network solver with computational fluid dynamics to investigate possible modifications to a commercial gas turbine combustor. MS Thesis, University of Pretoria .
[17] K. A. das Neves. (2018). Combustion analysis on a CFM56 engine. Master's Degree in Aeronautic engineering, Beira Interior University.
[18] S.Farokhi. Aircraft Propulsion. second Edition, Wiley.
[19] A.Lefebvre-D.Ballal. (1979). Weak Extinction Limits of Turbulent Flowing Mixtures. Journal of Engineering for Power.
[20] E. zukoski. (1985). Aerothermodynamics of Aircraft Engine Components. AIAA.

نص كامل:

Journal of Simulation and Analysis of Novel Technologies in Mechanical Engineering

Research article

Conceptual design of the inlet flow preheat system for nozzle test equipment of a turbine engine

Behrooz Shahriari*, Hamid Farrokhfal, Mohammad Reza Nazari

Faculty of Mechanics, Malek Ashtar University of Technology, Iran

*shahriari@mut-es.ac.ir

(Manuscript Received --- 01 Oct. 2023; Revised --- 21 May 2024; Accepted --- 16 June 2024)

Abstract

Keywords: Turbofan engine nozzle, Nozzle ground test, Hot flow simulation, Combustion chamber design, Afterburner design.

1- Introduction

The testing of the main components of the turbine engines as well as the integrated testing of the entire propulsion system includes hours of ground and flight testing. The ground test of the main components, including the compressor, combustion chamber, turbine, afterburner, and nozzle, requires a special test room for each of these components. Compressed air supply facilities and the preheat section of the flow are needed to simulate the cold and hot inlet flow. Since the nozzle of the turbofan engine has an afterburner in addition to the main combustion chamber, the preheat system must have the ability to create the inlet temperature conditions for the nozzle in two modes, afterburner on and afterburner off. In fact, the preheat system used in the nozzle test room has two sections, the combustion chamber and the afterburner, which operate depending on the need. Therefore, the design of this ground preheat system has its own complexities, and the method of designing the main components of the nozzle and afterburner in air turbine engines can be used in this direction. However, due to the fact that this preheat system is on the ground, some air requirements such as minimum weight can be ignored or a cooling method other than air can be used to increase the cooling efficiency. The design of gas turbine combustion chamber and combustion analysis has been a serious challenge for researchers due to the existing complexities such as turbulence, chemical kinetics, thermal radiation and the production of pollutants. Large manufacturing companies use expensive programs to upgrade and improve the combustion chamber of turbine engines, and most of these methods have achieved their desired combustion chamber after initial designs with methods based on trial and error [¹]. Also, these companies have obtained their own rules by using laboratory tests and experience, which help in product development and improvement programs. These design rules provide solutions to achieve the desired chamber geometry given the inlet conditions. If a combustion chamber designer does not have access to these solutions, it is necessary to obtain his own design method using available sources such as articles and experiments. However, a significant number of empirical, semi-empirical, and analytical relationships are available that minimize the need for costly experimental tests. In addition, in the combustion chamber design process, the use of advanced computational methods and accurate modeling of combustion flow behavior can reduce the number of laboratory tests and make the design process shorter and less costly. Combustion chamber designers have always sought to design a combustion chamber that, in addition to high efficiency and low pressure drop, has an optimal outlet temperature and a long service life. Mattingly [²] and Lefebvre [³] presented a set of empirical, semi-empirical relationships and analytical models that are used in the design of conventional combustion chambers of turbine engines. They have also provided information and solutions regarding new generations of modern enclosures. Charest presented relationships for the preliminary design of a can-type combustor. The presented method was based on component modeling and phenomena such as droplet evaporation, heat transfer and jet mixing were modeled in it. He compared and evaluated the results of his method in which he used empirical and semi-empirical relations with the computational fluid dynamics solution of a combustion chamber [⁴]. In another study, Stuttsford and Rubini modeled flow and heat transfer using a network method. This method divides the chamber into several separate and internally connected control volumes. Then the continuity, momentum and energy equations are discretized for different control volumes and by solving the equations, internal flow and temperature variables are obtained at different points [⁵]. Khandelwal designed the combustion chamber with the approach of reducing pollutants and studied the effect of different geometrical parameters on the performance of the combustion chamber composite diffuser [⁶]. In the process of designing a new generation of Rolls-Royce combustion chambers, Wankhede used a combined method using CFD and experimental relationships [⁷]. Jai-Houng Leu designed and simulated a combustion chamber used in a gas turbine, the fuel used is of low heating value type, and the source of this fuel is the gaseous cation produced in the combustion of waste plastic. He also evaluated the performance of the designed chamber according to previous works and one-dimensional flow characteristics [⁸]. Yize Liu et al. developed a general framework for the preliminary design of air turbine combustion chambers. Design elements were a combination of flow distribution, combustion chamber size, heat transfer and cooling, pollutant levels, and other performance parameters. The presented numerical method was based on experimental and semi-experimental methods and the accuracy of the design was checked using previous works [⁹]. The review of the research conducted in the field of combustion chamber and afterburner of turbine engines show that so far no research has been presented on the design and analysis of the two-stage engine preheat system. In the present study, a combustion chamber with an afterburner (without the use of a turbine) was used in an innovative way to supply air to the nozzle test system. At first, according to the semi-experimental and analytical relations, the geometric and thermodynamic characteristics of the combustion chamber were determined, and then the obtained results were validated with the data of a similar combustion chamber. Finally, the temperature of the exit stream from the combustion chamber has been increased to the desired value by an afterburner designed by reverse engineering method and empirical relationships.

2- Combustion chamber design algorithm and method

In this section, the combustion chamber is done at the conceptual design level. In the first step, the diameter of the liner and the reference surface of the combustion chamber are determined. Then, according to the inputs of the problem and the relationships extracted from the authoritative sources, the two-dimensional geometry of the combustion chamber is produced and by using the relations of the combustion flows, the characteristics of the output flow of the combustion chamber are predicted. Fig. 1 shows the combustion chamber design algorithm.

combustion chamber2

Fig. 1 Combustion chamber design algorithm

2-1- Reference diameter and liner diameter

The first step in designing the combustion chamber is to determine the reference diameter . The reference diameter is the same as the diameter of the combustion chamber shell, and in most references, the reference diameter is calculated by considering chemical considerations or pressure drop. After determining the reference diameter, the diameter of the liner is calculated. Fig. 2 shows the reference diameter and liner diameter of conventional

combustion chambers. Estimation of the reference diameter can be done based on different methods. Various algorithms and methods have been presented in different references to estimate the reference diameter and liner diameter and how these two sections are related. In this work, four different methods are used to obtain these values as follows.

One of the reference diameter estimation methods is the aerodynamic requirements method, which is based on the pressure drop ratio and pressure factor.

$C:\Users\Rose Computer\Desktop\afterburner\55.png$

Fig. 2 Reference diameter and liner diameter in combustion chamber types

In general, if the combustion chamber is large enough to accommodate a certain pressure drop, it will be large enough to accommodate the chemical reactions. Sometimes, depending on the conditions, the diameter obtained from the reaction rate equations may be greater than the diameter obtained from the aerodynamic method, in which case the combustion engineer is responsible for determining the correct method [¹⁰]. In this method, the reference level is calculated from equation (1).

(1)

where is called the total pressure drop and expresses the ratio of the total pressure drop to the total inlet pressure at the time of ignition. This parameter is usually expressed as a percentage (usually between 4 and 8) and depends on the operational conditions of the combustion chamber. The term is called the pressure factor and it expresses the ratio of the total pressure to the reference dynamic pressure at the time of ignition. This parameter is very important for combustion engineers, because it represents the resistance of the current applied between the compressor outlet and the turbine inlet. From aerodynamic point of view, this parameter can be considered equivalent to the Drag coefficient, which is one of the constant characteristics of the combustion chamber (it does not depend on the operating conditions). Also, this parameter indicates the total pressure drop in the diffuser and throughout the liner, which is shown as Eq. (2) [3].

(2)

In addition to aerodynamic requirements, the combustion chamber must be designed in such a way that it is possible to carry out combustion processes in it. The method of combustion requirements, taking into account the geometrical conditions necessary for combustion, determines the reference diameter necessary to carry out the combustion reaction in different working conditions. Lefebvre stated that for any specific ratio of fuel to air, the combustion efficiency (η) is a function of the parameter θ, which is equal to Eq. (4).

	(3)
	(4)

In the above equation, parameter b is the temperature correction and its value can be considered as 300 degrees Kelvin on average. In order to calculate more accurately, the temperature correction parameter as a function of equivalence in the initial area of the combustion chamber can be expressed as equation (5).[2]


	(5)

where b is in Kelvin and is the equivalence ratio of fuel and air mixing in the initial area of the combustion chamber. Another method that calculates the reference diameter by considering the combustion requirements and flame stability is called the combustion loading method, in which the volume of the combustion chamber is calculated with loading parameters. The study [¹¹] presented Eq. (6) for the loading parameter:

(6)

where, is the mass flow rate in Kg/s, V is the combustion chamber volume in , is the inlet pressure in atm and is the inlet temperature in Kelvin. The maximum loading parameter value occurs for the highest flight altitude, the lowest flight Mach Number and the coldest weather conditions.

In this situation, in order to create a sufficient margin of confidence to prevent flame extinction and achieve acceptable efficiency, this parameter is considered less than 50. Also, for the conditions of the sea level and the highest thrust, the loading parameter is recommended to be less than 10 and preferably less than 5. The combustion intensity parameter is a measure of the heat released per unit volume of the combustion chamber and is defined as Eq. (7)

(7)

where LHV is the minimum thermal energy of the fuel in terms of (KJ/Kg) and is the combustion efficiency in the combustion chamber. From another point of view, the loading parameter is also a measure of the difficulty of combustion, so it is desirable to have a low value. For design conditions (operating conditions at sea level and maximum thrust), the combustion intensity should be less than 60. This value is achievable for industrial engines, but it can be challenging for aero engines. Also, in the industrial gas turbine, this value may be considered lower due to the larger available combustion space and due to the possibility of using a converter and the need for less heat. The volume of the combustion chamber should be such as to ensure that the loading parameter corresponds to the intensity of combustion. The volume of the chamber can be calculated from Eq. (8):

(8)

Where is the ratio of thermal coefficient at constant pressure to constant volume and is the ratio of total pressure to static pressure. Having the cross section of the liner and the volume of the combustion chamber, the length of the chamber is obtained from Eq. (9):

(9)

Now the residence time can be checked with Eq. (10). The residence time is the time it takes for an air molecule to pass through the combustion chamber, which must be more than 3 ms in order to fully mix fuel and air and have a complete combustion. It is important to note that by increasing the mass flow rate entering a combustion chamber with a certain geometry, the flow Mach Number increases and as a result, the residence time decreases.

(10)

Different sources have provided several methods to estimate the area of the combustion zone .In a simple relation, the ratio of can be considered for the can and annular chambers and the ratio of for the annular can combustion chamber [11]. Also, Eq. (11) provided by Bragg:

(11)

In addition, in the reference [2], in order to maximize the penetration ratio of the cooling fluid jet, Eq. (12) is proposed:

(12)

Because Eq. (12) involves more parameters in the design, it is used in this study.

2-2-Diffuser and snout

After leaving the compressor, the air flow enters the diffuser of the combustion chamber, whose main function is to reduce the Mach (velocity) of the incoming flow and increase the static pressure. In the process of designing the diffuser, first the efficiency and then its dimensions are determined. The efficiency of the diffuser is a function of its divergence angle and is defined as follows:

(13)

Then the flow weight function is calculated by Eq. (14).

(14)

The flow weight function can be calculated at the inlet by knowing the Mach number at station 32. Also, the ratio of total to static pressure is obtained using Eq. (15). Although, the total pressure at station 32 is not known, but it is a function of the ideal pressure recovery coefficient and can be calculated using Eqs. (15) to (20) [¹²]:

	(15)
	(16)
	(17)
	(18)
	(19)
	(20)

Then, the flow function is calculated again and if the new value matches the previous value, the solution is converged, otherwise, the solution method is repeated with the new assumption for the ideal recovery coefficient. The snout is placed between the diffuser and the initial area of the combustion chamber (before the swirler). The snout of the combustion chamber should be designed in such a way that the amount of air desired by the designer enters the annular section and the rest of the flow enters the combustion chamber through the holes of different areas. Therefore, the most important task of the snout is to divide the flow between the inlet core and the annular area of the combustion chamber. The cross-sectional area of the entrance snout is calculated from Eq. (21) [2].

(21)

where is the snout coefficient, which in the ideal case (uniform output assumption) is considered to be equal to unity. Also, the mass flow rate of air passing through the snout is considered to be approximately equal to . Eq. (22) is used to calculate the pressure drop in the snout section. [12]

(22)

2-3- Swirler

One of the components of the combustion chamber that has a direct effect on the flame structure, the level of pollutants and the overall efficiency of combustion is the swirler. The main task of the swirler is to convert a part of the axial velocity component into radial velocity [10]. Another task of the air swirler and primary holes is to create a return and vortex flow. Strong and stable return currents in the primary area cause flame stability and good fuel-air mixing, which makes the combustion chamber perform well in a wide range of flight conditions. The usual type of swirler includes a number of blades (usually 8 to 10) that are placed around the fuel injector at a specific angle [12]. The pressure drop in the combustion chamber includes the total pressure drop in the swirler, snout and diffuser. The swirler pressure drop is calculated from Eq. (23).

(23)

The basis of the design of the dimensions of the impeller is based on the amount of pressure drop obtained in the thermodynamic analysis (zero dimension) of the gas turbine cycle. knight and walker proposed the following equation to express the relationship between pressure drop and the driving surface [¹³]:

(24)

where is the flow rate of the air entering the swirler, which is considered to be about 30-70% of the air in the primary entrance area. Also, based on experimental results, the amount of this air is considered to be 3-12% of the total incoming air (source 23). is also the air rotation coefficient, which is considered to be 1.3 in straight blades and 1.15 in curved blades. The swirler are usually made of straight blade type and the air circulation angle( )in them is considered to be about 45 to 70 degrees. After calculating the surface of the swirler, the outer radius of the swirler is calculated from Eq. (25).

(25)

Finally, the amount of rotation that the swirler applies to the flow is expressed by the swirler number, which is calculated by Eq. (26):

(26)

2-4-Primary, secondary and dilution area

After injection, the fuel passes through three different areas. The first zone is the primary combustion zone, the task of this zone is to maintain the flame and also to provide enough time, temperature, and agitation so that the fuel sprayed in this zone completely combines with the incoming air, and the chemical energy hidden in it is converted into thermal energy. In other words, provide the conditions for a complete combustion. Fig. 3 shows the schematic of the rotating area of the combustion chamber.

Fig. 3 Schematic of circulation area [12]

In order to place ignitors, the concept of magic circle is used. As shown in Figure 6, the magic circles are surrounded by the midline, the liner wall, and the dome wall of the liner. Also, this area is rich in fuel and air mixture, and the diameter of the magic circles is considered to be equal to half the diameter of the liner. The distance between the ignitors is calculated from Eq. (27) [¹⁴]:

$C:\Users\Rose Computer\Desktop\afterburner\45.PNG$

Fig. 4 Schematic of the rotating area and magic circles [¹⁵]

(27)

where is the divergence angle of the diffuser, whose value is considered to be 60 degrees in conventional combustion chambers. Also, the length of the circulation area is equal to:

(28)

Finally, knowing the swirler number and diameter of the swirler, the length of the primary area can be calculated with the following equation [2]:

(29)

In the secondary area, with the addition of air currents, the combustion process is complete and we have the highest temperature (among the different areas of the combustion chamber), therefore, wall cooling in the middle area is one of the most important issues in the design of the combustion chambers.

Then, in the dilution zone, the combustion flow is combined with the incoming air from the dilution zone and the output flow is regulated according to the turbine blades. Fig. 5 shows the distribution of air in different areas of the combustion chamber.

Fig. 5 Schematic of air distribution in different areas of the combustion chamber [2]

In this section, the details of calculating the holes of these two areas are briefly stated based on the ratio of the maximum penetration depth to the required jet diameter. The ratio of the maximum penetration depth to the jet diameter () is defined by Eq. (30) [3]:

(30)

where θ is the penetration angle of the fluid jet and J is the ratio of the momentum flux of the fluid jet entering the gas (transverse flow), which is expressed as equation (31):

[1] References

[] W. Dodds and D. Bahr. (1990). Combustion system design. Design of modern gas turbine combustors, pp.343-476.

[2] [] J. D. Mattingly. (2002). Aircraft engine design: AIAA.

[3] [] A. H. Lefebvre. (2010). Gas turbine combustion. CRC Press.

[4] [] M. R. J. Charest. (2006). Design methodology for a lean premixed prevaporized can combustor. MS Thesis, Library and Archives Canada,Carleton University, Department of Mechanical and Aerospace Engineering.

[5] [] P. J. Stuttaford and P. A. Rubini. (1996). Preliminary gas turbine combustor design using a network approach. In ASME International Gas Turbine and Aeroengine Congress and, Exhibition.

[6] [] P. J. Stuttaford and P. A. Rubini. (1996) Preliminary gas turbine combustor design using a network approach. in ASME International Gas Turbine and Aeroengine Congress and, Exhibition.

[7] [] M. J. Wankhede. (2012). Multi-fidelity strategies for lean burn combustor design. PhD Thesis, University of Southampton.

[8] [] Jai-Houng Leu. (2010). Design and simulation validation of LHV fuel combustor. Journal of Information and Optimization Sciences, 31:6,1321-1336.

[9] [] Yize Liu, Xiaoxiao Sun, Vishal Sethi, Yi-Guang Li, Devaiah Nalianda, David Abbott, Pierre Gauthier,Bairong Xiao and Lu Wang. (2021). Development and application of a preliminary design methodology for modern low emissions aero combustors. Journal of power and energy. Vol. 235(4) 783-806.

[10] [] J. W. Sawyer. (1966). Gas Turbine Engineering Handbook: Editor. John W. Sawyer, vol. , Gas Turbine Publications.

[11] [] P. P. Walsh and P. Fletcher. (2004). Gas turbine performance. John Wiley & Sons.

[12] [] R. Rezvani. (2010). A Conceptual Methodology for the Prediction of Engine Emissions. PhD Thesis, Georgia Tech. University.

[13] [] H. Knight and R. Walker. (1953). The component pressure losses in combustion chambers. Gt. Brit. National Gas Turbine Establishment, Farnborough, Hants, England.

[14] [] M. R. J. Charest. (2006). Design methodology for a lean premixed prevaporized can combustor. MSThesis, Library and Archives Canada Bibliothèque et Archives Canada, Carleton University,
Department of Mechanical and Aerospace Engineering.

[15] [] A. Costa Conrado, P. T. Lacava, A. C. P. Filaho, M. D. S. Sanches. (2004). Basic design principles for gas turbine combustor. Proceedings of the 10o Brazilian Congress of Thermal Sciences and Engineering.

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Conceptual design of the inlet flow preheat system to the nozzle test equipment of a turbine engine

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