Feasible Concept of an Air-Driven Fan with a Tip Turbine for a High-Bypass Propulsion System
energies
Article Feasible Concept of an Air-Driven Fan with a Tip Turbine for a High-Bypass Propulsion System
Guoping Huang, Xin Xiang *, Chen Xia, Weiyu Lu and Lei Li College of Energy and Power, Nanjing University of Aeronautics and Astronautics, Nanjing 210016, Jiangsu, China; [email protected] (G.H.); [email protected] (C.X.); [email protected] (W.L.); [email protected] (L.L.) * Correspondence: [email protected]; Tel.: +86-139-2144-7677
Received: 19 October 2018; Accepted: 27 November 2018; Published: 30 November 2018
Abstract: The reduction in specific fuel consumption (SFC) is crucial for small/mid-size cost-controllable aircraft, which is very conducive to reducing cost and carbon dioxide emissions. To decrease the SFC, increasing the bypass ratio (BPR) is an important way. Conventional high-BPR engines have several limitations, especially the conflicting spool-speed requirements of a fan and a low-pressure turbine. This research proposes an air-driven fan with a tip turbine (ADFTT) as a potential device for a high-bypass propulsion system. Moreover, a possible application of this ADFTT is introduced. Thermodynamic analysis results show that an ADFTT can improve thrust from a prototype turbofan. As a demonstration, we selected a typical small-thrust turbofan as the prototype and applied the ADFTT concept to improve this model. Three-dimensional flow fields were numerically simulated through a Reynolds averaged Navier-Stokes (RANS)-based computational fluid dynamics (CFD) method. The performance of this ADFTT has the possibility of amplifying the BPR more than four times and increasing the thrust by approximately 84% in comparison with the prototype turbofan.
Keywords: high-bypass ratio engines; tip turbine; air-driven fan; aerodynamic analysis
1. Introduction In recent years, small/mid-size cost-controllable aircrafts, such as unmanned aerial vehicles (UAVs), small airplanes for general aviation and cruising missiles, have demonstrated considerable value in the military/civilian fields [1,2]. The propulsion systems of these aircraft have two important requirements, namely, long voyage and low cost of use and manufacturing. As an important performance parameter, specific fuel consumption (SFC) is related to the two requirements. Several traditional techniques, such as increasing the overall pressure ratio (OPR), turbine inlet temperature and components efficiency, can be performed to improve the SFC. However, after an extensive development, these parameters have faced recent technology limitations, and raising the core thermal efficiency for conventional turbofan engines is becoming increasingly disadvantageous. In particular, increasing bypass ratio (BPR) is an appropriate method. Zimbrich in [3] suggests that, ideally, the SFC has a positive trend with increasing BPR, whereas the optimum fan pressure ratio has a continuous downward trend. Many segments of the aircraft industry and NASA (National Aeronautics and Space Administration) anticipate the application of high-bypass propulsion systems to small and fast transport aircraft [4,5]. Therefore, a high-BPR turbofan engine will face the spool-speed conflict with a fan and a low-pressure turbine (LPT). For a given core size, a high BPR indicates an increasing fan diameter. On the one hand, considering the tip speed for an acceptable level of a fan’s shock loss and the permissible centrifugal stress at a fan blade root, the rotational speed of a fan spool is limited at a low
Energies 2018, 11, 3350; doi:10.3390/en11123350 www.mdpi.com/journal/energies Energies 2018, 11, 3350 2 of 16 level. On the other hand, the LPT that drives the fan requires a high rotational speed to reduce stage loading and remain working at a high efficiency. This conflict will become increasingly aggressive with the increase in the BPR [6]. Several size problems will occur when the fan diameter is limited if the BPR is increased by scaling down the core engine size. For example, the blades of a high-pressure system (including the compressor and turbine) are too short, and the tip leakage losses consequently increase, thereby finally reducing the efficiency. Several kinds of turbofan, such as a tri-spool turbofan and an aft-fan machine, have been developed for a high-BPR engine [7,8]. These kinds of turbofan can only relieve the conflict. However, when the BPR rises to ultra-high, the effects of the techniques in those engines are limited. The geared turbofan engine (GTF) is a successful technology for high/ultra-high BPR engines in commercial airplanes [9]. By using a gearbox, the GTF can enable the fan and LPT to rotate at different rotational speeds. The cost of developing the gearbox and improving the core spool is relatively small when compared with their benefits to commercial airplanes. In small/mid-size cost-controllable aircraft, applying the GTF may be another situation. The factors, including the limitation of a spool size, the cost of such a sophisticated geared-box, and the complicated rotor dynamics problems, will weaken the advantages of gearbox in the small/mid-size cost-controllable aircraft. In fact, the GTF has been used in several mid-size thrust engines, such as the TFE731 and ALF502, in the early 1970s. However, GTF is not extensively used at that time considering the cost and maintenance. Tip turbine is another concept for solving the conflict and has been adopted in very high BPR turbofan engines. The mounting of the turbine at the tip of a fan blade can enable the fan and turbine to work at their optimal stage loading. In addition, the mechanical connection between the fan and the core rotor system is released through a gas-driven method, thus avoiding the changes in the core engine, especially the rotor spool. In the 1960s, the tip turbine has been used in lift fans for a vertical take-off and landing (VTOL) aircraft. NASA has investigated such a concept used in XV-5A and other further developed aircraft [10,11]. For the VTOL aircraft, the BPR of its engine must be high under a take-off condition to ensure the thrust against its weight. Reference [10] demonstrates that such a tip turbine cruise and lift fans can reach a BPR range of 10–100. NASA has conducted considerable research on such a kind of concept and has tested its turbine efficiency [12,13]. In the 2000s, researchers at Cranfield University studied a tip turbine-driven propulsion fan concept which has a potential future high-BPR turbofans [14]. Hatta [15] proposed a development study of an ATREX (Air-Turbo-Ram Expander cycle) engine, an air turbo ramjet with an expander cycle, which is a turbojet engine at a subsonic to Mach 2 flight. The ATREX uses a tip turbine configuration to provide compactness and light weight of the turbomachinery. According to existing public literature, a tip turbine functions at a high temperature and contains dozens to hundreds of blades. However, the heat conduction from a tip turbine to its fan will make maintaining a high efficiency difficult for a fan. An excessive number of blades will also impose additional stringent requirements on the strength of the leaf roots, especially the turbine blade at the tip. The reduction in the blade number of the tip turbine and its air temperature while ensuring driving capability will cause advantages to its use. The present research proposed an aerodynamic consideration of an air-driven fan with a tip turbine (ADFTT) as a potential solution for a high-bypass propulsion system. The tip turbine had the same number of blades with its fan and works with bypass air from a core engine. The aerodynamic and thermodynamic analyses of such an ADFTT are conducted to show its potential in increasing BPR. As a demonstration, a typical small-thrust turbofan was selected as the prototype and applied the ADFTT to improve this model. The performances of the ADFTT were numerically simulated through the RANS based CFD method. Energies 2018, 11, 3350 3 of 16
2.Energies Principle 2018, of11, thex FOR ADFTT PEER REVIEW and Its Application to the High-Bypass Propulsion System 3 of 16 2.1. Principle and Brief Structure of the ADFTT The ADFTT is an aerodynamic component with two parts (Figure 1), namely, the fan and the tip The turbine ADFTT mounted is an aerodynamic at the tip of component the fan. The with tip two turbine parts can (Figure achieve1), namely, a positive the fan circumferential and the tip turbinevelocity mounted at a low at the spool tip of speed. the fan. The The two tip turbine parts are can attached achieve a to positive each othercircumferential in a structural velocity and at aaerodynamic low spool speed. design: The thetwo total parts pressure are attached at the to each outlet other of inthe a fanstructural and the and turbine aerodynamic must be design: equal the(minimizing total pressure the atmixing the outlet loss); of the the blade fan and numbers the turbine of the must fan and be equal the turbine (minimizing rotor must the mixing be the loss);same the(considering blade numbers the blade of the strength)fan and the and, turbine given rotor the must structure be the consistency same (considering and reliability, the blade the strength) turbine and,rotor given must the have structure the same consistency staking and and sweeping reliability, laws the turbinewith the rotor fan rotor. must have the same staking and sweepingThe lawsblade with of the the air fan-driven rotor. fan rotor can be joined through two methods. The first method uses a bladeThe bladeshroud of (Figure the air-driven 1, ①) fanthatrotor can divide can be the joined flow through paths twoof the methods. fan and Thethe firstturbine method, in order uses ato bladeavoid shroud the interferences. (Figure1, 1 The) that second can divide method the uses flow a paths blade of shoulder the fan and (Figure the turbine, 1, ②) to in join order the to fan avoid and theturbine interferences. blades. The The blade second shoulder method is useslighter a blade than shoulderthe blade (Figureshroud1 but, 2 presents) to join themore fan difficulties and turbine for blades.its design. The bladeFor the shoulder fan, the is number lighter thanof the the blade blade will shroud be limited but presents because more the difficulties small number for its of design. blades Forcan the reduce fan,the the number centrifugal of the force blade on willa fan be disk limited and can because lead theto a small small number wetted ofarea blades and minimal can reduce friction the centrifugalloss. force on a fan disk and can lead to a small wetted area and minimal friction loss.
Tip-Turbine Rotor ① Blade Shroud OR ② Blade Shoulder
Fan Rotor
R o D t i at re in ct g io n
FigureFigure 1. 1.3D 3D Model Model of of the the ADFTT ADFTT rotor. rotor.
2.2.2.2 High-Bypass. High-Bypass Propulsion Propulsion System System that that Adopts Adopts the the ADFTT ADFTT TheThe ADFTT ADFTT can can eliminate eliminate the the mechanical mechanical limitations limitations of of a a fan fan (which (which produces produces the the optimum optimum thrust)thrust) from from the the core core engine engine and and alter alter the energy the energy transfer transfer method method aerodynamically. aerodynamically. The air for The driving air for thedriving ADFTT the can ADFTT originate can originate from the from core engine,the core such engine, as a su turbojet,ch as a theturbojet, bypass the of bypass turbofan of andturbofan the mixedand the nozzle mixed of the nozzle turbofan. of the Figure turbofan.2 illustrates Figure a2 preliminary illustrates a high-bypass preliminary propulsion high-bypass system propulsion that adaptssystem with that the adapts ADFTT. with The the system ADFTT. mainly The system includes mainly five components, includes five namely, components, (1) an air-drivennamely, (1) fan an rotorair-driven (including fan rotor tip turbine), (including (2) tip an turbine), additional (2) stator; an additional (3) an air stator; bleeding (3) an channel, air bleeding (4) an channel, additional (4) casingan additional and (5) a casing rotational and casing. (5) a rotational The air of thecasing. bypass The of air a core of the turbofan bypass is oftransferred a core turbofan to the tip is turbinetransferred (mounted to the at tip the turbine tip of a (mounted fan blade) at through the tip the of air-bleedinga fan blade) channel.through Similarthe air- tobleeding the GTF channel. which transfersSimilar to the the energy GTF which from thetransfers LPT to the the energy airflow from in the the bypass, LPT to the ADFTTairflow in can the also bypass, transfer the energyADFTT fromcan thealso bypass transfer of energy a core enginefrom the to anbypass increasing of a core airflow engine in the to additionalan increasing bypass. airflow The in ADFTT the additional rotates atbypass. a low speed The ADFTT and works rotates as an at ‘aerodynamic a low speed geared and works box’. asBy an using ‘aerodynamic such an aerodynamic geared box’. geared By using box, thesuch mechanical an aerodynamic limitations geared of the box,fan (which the mechanical produces the limitations optimum of thrust) the fan from (which the core produces engine can the beoptimum removed, thrust) and the fromenergy the is transferred core engine aerodynamically. can be removed, Such a designand the concept energy is the is subject transferred of a Chineseaerodynamically. invention patentSuch a [desi16]. gn concept is the subject of a Chinese invention patent [16].
Energies 2018, 11, 3350 4 of 16 Energies 2018, 11, x FOR PEER REVIEW 4 of 16
Figure 2 2.. Structure sketch of the ADFTT applied to the prototype turbofan.turbofan.
3. Theoretical Thermodynamic Analysis of the ADFTT
3.1.3.1. Basic Equilibrium Principles As mentioned previously, the ADFTT acts as an aerodynamic box, whose basic function is to transfer energy to the airflow in additional bypass, and we defined its efficiency as ηADFTT. To design transfer energy to the airflow in additional bypass, and we defined its efficiency as ADFTT . To this ‘aerodynamic gearbox’, the working procedures of the device should be analysed. Several basic design this ‘aerodynamic gearbox’, the working procedures of the device should be analysed. equilibrium equations related to overall parameters can be listed as follows: Several basic equilibrium equations related to overall parameters can be listed as follows: The energy transforming procedure is put forward at first place. The energy contained in the The energy transforming procedure is put forward at first place. The energy contained in the additional bypass originates from the bypass of the core engine, although the mass flow rate of the additional bypass originates from the bypass of the core engine, although the mass flow rate of the additional bypass is large. The energy equilibrium formula can be written as: additional bypass is large. The energy equilibrium formula can be written as: − − k 1 kk11k 1 E = η ·E = η ·m ·C ·T∗ π k − = m ·C ·T ∗ π k − ADFTT ADFTT B ADFTT bypass p 1 **f 0 kk1 ADFTT p 1 ADFTT 1 , (1) EADFTT ADFTT E B ADFTT m bypass C p T f0 11 m ADFTT C p T ADFTT , (1) 11 where EADFTT is the energy that transformed by ADFTT (from tip turbine to air-driven fan), mADFTT where E is the energy that transformed by ADFTT (from tip turbine to air-driven fan), is the massflowADFTT rate of the ADFTT (including air-driven fan and tip turbine), Cp is the specific heat m ∗ C capacityADFTT is of the air, massflowT1 is the total rate temperature of the ADFTT of air (including of the inlet air which-driven equals fan and to the tip ambient turbine), temperature, p is the * k sπpecificADFTT heatis the capacity pressure of ratio air, ofT1 ADFTT is the (after total the temperature mixing of of air-driven air of the fan inlet and which tip-turbine), equals tois the specific heat ratio of air, ηADFTT is the equivalent isentropic efficiency of ADFTT, EB is the energy ambient temperature, ADFTT is the pressure ratio of ADFTT (after the mixing of air-driven fan and contains in the bypass flow of core engine, mbypass is the massflow rate of the bypass flow of core-engine, tip-turbine), k is the specific heat ratio of air, ADFTT is the equivalent isentropic efficiency of π f 0 is the pressure ratio of the fan of core engine. ADFTT, EB is the energy contains in the bypass flow of core engine, mbypass is the massflow rate The relationship between mADFTT and mbypass can be expressed as: of the bypass flow of core-engine, f 0 is the pressure ratio of the fan of core engine. m mADFTTm B The relationship between BADFTTADFTT =and bypass =can be expressed·mADFTT as.: (2) mcore mbypass m B BmADFTT . where the B is the ratio of the massflowADFTT in additional bypass ADFTT divided by the massflow of the(2) ADFTT mmcore bypass high pressure system of core engine, mcore is the massflow of the high pressure system of core engine, whereB is the the bypass BADFTT ratio is of the core ratio engine. of the massflow in additional bypass divided by the massflow of the high The pressure power system extracted of fromcore engine,the tip turbine mcore is and the the massflow power consumed of the high by the pressure fan must system be matching. of core Theengine, formula B is can the be bypass written ratio as: of core engine. The power extracted from the tip turbine and the power consumed by the fan must be matching. ∗ 1 ∗ k−1 ( − ) = ( k − ) The formula can bem writtenTTCpTf 0as1: k−1 ηTT mADFCpT1 πADF 1 /ηADF, (3) πTT k 1 k 1 m C T**(1 ) m C T ( k 1) / , TT p f01k 1 TT ADF p ADF ADF (3) k TT
Energies 2018, 11, 3350 5 of 16
where mTT is the massflow rate of tip turbine which is equals to the mbypass, πTT is the pressure drop ∗ ratio of tip turbine, ηTT is the efficiency of tip-turbine, T1 is the total temperature of the inlet air of tip turbine, mADF is the massflow rate of air-driven fan, ηADF is the isentropic efficiency of the air-driven fan. These equations are the basic equilibrium equations of the ADFTT. Additional results can be deduced from these equations. For example, to minimise the mixing loss in the additional bypass nozzle, the total pressure of the fan and turbine outlets must be the same. Thus, the pressure ratios of the fan of core engine, the air-driven fan and tip turbine has a relationship as:
π f 0 πADF = = πADFTT, (4) πTT
With the help of these equations, we could calculate the overall parameters of ADFTT and get the performance of ADFTT.
3.2. Thermodynamic Analysis of the ADFTT Effects A thermodynamic analysis is constantly the most important part that can show the loss and gain of a thermodynamic performance. An additional thermodynamic cycle to transfer energy will cost energy loss. We define the transfer efficiency for each turbofan to judge the amount of energy that the bypass airflow generates from the LPT. By ignoring the heat dissipation, we can easily deduce certain results from energy equations. For a normal turbofan, the energy transfer efficiency is:
ηt = ηmη f 0ηo, (5) where ηm is the mechanical efficiency of rotating shaft, ηo is the efficiency of the outlet nozzle. For the GTF, the energy transfer efficiency is:
ηGTF = ηmη f 0ηoηg, (6) where ηg is the mechanical efficiency of the geared box. The ADFTT can be deduced as:
k−1 k−1 k−1 k−1 k−1 k − k k − k k 1 π f 0 πADF π f 0 πADF π − 1 η f 0 η = ( η η + η + ADF )η η , (7) ADFTT k−1 k−1 TT f 0 k−1 TT k−1 η ch mix k − k k k − ADF π f 0 1 π f 0 π f 0 π f 0 1 where ηmix refers to the equivalent efficiency due to the mix loss in the bypass, and ηch refers to the equivalent efficiency of the air-bleeding channel. This equation shows that the efficiency of the ADFTT is related to not only the components’ efficiencies but also the pressure ratios (they could be deduced from other top-level parameters). After the transfer procedure, the energy is redistributed into the airflow in the additional bypass with a large mass flow, and then thrust will be increased. The thrust of the prototype is:
Fcore = Fcore−nozzle + FB = (1 + α)FB, (8) where Fcore is the thrust of core engine, Fcore−nozzle is the thrust produced by the nozzle of high pressure system of core engine, FB is the thrust produced by the bypass of core engine (assume the core engine is separate exhaust), α refers to the ratio of the nozzle thrust divided by the bypass thrust. Energies 2018, 11, 3350 6 of 16
Considering the ideal situation and ignoring several small losses (e.g., the friction loss), the energy contained in the bypass is: 2 2 mBVB FB EB = = , (9) 2η f 0 2mBη f 0 where VB is the airflow velocity of the outlet of bypass. This part of the energy is redistributed by the ADFTT. Then, the thrust of the ADFTT can be expressed as:
!1/2 0 ηADFTTηADF BADFTT FADFTT = Fcore−nozzle + FB = α + FB. (10) η f 0B
0 where FB is the thrust produced by additional bypass. Thus, the thrust is amplified:
1/2 1/2 + ηADFTT ηADF BADFTT ηADFTT ηADF BADFTT − F α η B η B 1 ADFTT = f 0 = 1 + f 0 (11) Fcore 1 + α 1 + α
Therefore, the ADFTT can increase the thrust by redistributing energy into an additional airflow, although the process consumes certain energy. In particular, the losses related to the fan, air-bleeding channel, tip turbine and air-driven fan are important. The thrust will increase with the BPR of the ADFTT (BADFTT) and will be influenced by an equivalent efficiency (ηADFFF). The FADFTT calculated from Equation 11 is in ideal situation and ignored the losses in nozzle of additional bypass. Usually, the thrust will be decrease to approximately 0.98 due to the loss in nozzle. The effect of the increase in thrust under a typical condition is demonstrated in Section 3.3.
3.3. Engine Performance Estimation of the ADFTT The efficiencies of the components are all important (as mentioned in Section 3.2), that is, the loss in the air-bleeding channel and the efficiencies of the tip turbine and air-driven fan. In this preliminary research, these efficiencies are assumed by the recent aerodynamic design level of components. For example, the efficiency of the air-bleeding channel is estimated at 0.96 because the bleeding channel with a positive pressure gradient can have a high pressure recovery under a low Mach number condition (e.g., less than 0.3). The efficiencies of the tip turbine and air-driven fan are conservatively estimated at 0.86 and 0.87, respectively. Based on these efficiency estimations, ηADFTT is higher than 0.8. For example, a case has been studied on the basis of a prototype turbofan CF738 [17]. Several parameters at a design point are set to π f 0 = 1.7, πADF = 1.09, ηTT = 0.86, η f 0 = 0.91, ηADF = 0.87, ηch = 0.96 and ηADFTT = 0.807 in accordance with Equations mentioned above. Moreover, the BPR can be increased from 5.3 to 25 or higher (see it in Table1). The efficiency of the ‘aerodynamic geared box’ seems unimpressive, the effect of amplifying the BPR is crucial and will reduce the SFC effectively. This result is similar with reference [18]: although the thermodynamic efficiency of geared fan is better than the tip turbine, the tip turbine arrangement remains a light and efficient system. By applying the ADFTT concept, an exploratory design is studied on the basis of a prototype turbofan CFE738. Several top-level parameters are listed in Table1. Several components’ efficiencies at a design point are set as it is mentioned above. The design target is determined to ensure that the BPR is at approximately 27.0. The thrust is amplified by 1.89. As an exploratory design, the ADFTT remains its potential for further improvement on efficiencies of components (especially tip turbine and fan). Therefore, its ability to increase the bypass ratio is reflected, which is the main contribution to the thrust amplification. Energies 2018, 11, 3350 7 of 16
Table 1. Parameters of the prototype turbofan and ADFTT.
Parameters CFE738 Preliminary ADFTT BPR 5.3 27.8 pressure ratio of bypass 1.7 1.085 Mass flow rate (kg/s) 95.3 422.0 Thrust (kN) 25.5 48.2
FADFTT/Fcore - 1.89 ηch - 0.96 ηTT - 0.86 ηADF - 0.87 ηADFTT - 0.807
4. Exploratory Design of the ADFTT
4.1. Top-Level Parameters of the Exploratory Design To demonstrate the ADFTT effects, an exploratory design is studied on the basis of the prototype turbofan of a general flight engine. The first step for exploratory design is to determine the top-level parameters of the propulsion system. These parameters determine the performance of the propulsion system and are used to guide the design of a component. Based on the theories in Section3, we can estimate the performance of a turbofan and ADFTT approximately if several assumptions are introduced as follows: (1) The core engine remains at the same performance (same thrust, same fuel consumption). (2) All the energy transferred to the air of the (additional) bypass is converted into the kinetic energy of the air. (3) Ignoring the nozzle loss to simplify the calculation. From these top-level parameters, we can calculate the design parameters of each component, such as mass flow rate, pressure ratio/drop and efficiency. In fact, the equivalent efficiency of the ADFTT is crucial, and ensuring the efficiencies of the aerodynamic components to be as high as possible, especially for the tip turbine and fan, is necessary.
4.2. Profiles Design of the ADFTT Blades The design processes of the fan and tip turbine, except the matching design process, are the same as the other turbomachinery; these design processes include preliminary, throughflow, 2D blading and 3D blading designs [19]. Notably, considering the structure consistency and reliability, the turbine rotor must have the same staking and sweeping laws with a fan rotor. The blade of the air-driven fan rotor can be joined through two methods. The first method uses a blade shroud (Figure1, 1 ) which can divide the flow paths of the fan and turbine and avoid the interferences. The second method uses a blade shoulder (Figure1, 2 ) which is lighter, but more challenging to design. In accordance with the top-level parameters listed in Table1, the aerodynamic parameters of the fan and turbine can be calculated predictively. Then, a geometry size and preliminary blade profile can be built. In this research, the ADFTT rotor, the most important and complex part, is completely built in the method 1 . The first stator of the tip turbine is built to ensure the integrity of tip turbine. Since the tip turbine mounted at the tip of the fan, the two parts are attached to each other in a structural and aerodynamic design. Several matching methods must be emphasised as follows: The blade numbers of fan and turbine rotor must be the same (considering the blade strength); the total pressure at the outlet of the fan and turbine must be equal (minimizing the mixing loss); the relative flow directions in each path should be similar (minimizing the shearing loss). The profiles of the ADFTT are depicted in Figure3 (they are the final profiles of the blades at present). In this exploratory design, we try not to restrict the design method of ADFTT or structure of blade, including the skewed and swept law. However, the fan has a skewed and swept blade to adapt the diameter changes from hub to tip. The rotor of the tip turbine is also a swept blade, whose swept Energies 2018, 11, x FOR PEER REVIEW 8 of 16
The profiles of the ADFTT are depicted in Figure 3 (they are the final profiles of the blades at present). In this exploratory design, we try not to restrict the design method of ADFTT or structure
Energiesof blade,2018 including, 11, 3350 the skewed and swept law. However, the fan has a skewed and swept blade8 of 16to adapt the diameter changes from hub to tip. The rotor of the tip turbine is also a swept blade, whose swept law is consistent from the fan. On the other side, the stator of the tip turbine can be a straight lawblade, is consistentand only slightly from the skewed fan. On considering the other side, the diameter the stator changes. of the tip All turbine the blades can be are a straightstaking by blade, the andcentre only of slightlygravity skewedand the considering3D structure the of tip diameter turbine changes. and air- Alldriven the bladesfan could are be staking seen byin Figure the centre 1. of gravity and the 3D structure of tip turbine and air-driven fan could be seen in Figure1. n n n o o o i i i t t t i i i s s s o o o P P P
l l l a a a i i i t t t n n n e e e r r r e e e f f f m m m u u u c c c r r r i i i C C C
Axial Position Axial Position Axial Position (a) Fan (b) Tip-Turbine Stator (c) Tip-Turbine Rotor
Figure 3.3. Section graphs of the ADFTT.ADFTT. (unit:(unit: mm;mm; 0% represents thethe hub;hub; 100%100% indicatesindicates thethe tip).tip).
For thethe fan,fan, the number of blades will be limited soso thatthat the centrifugal force ofof thethe fanfan diskdisk willwill bebe small. small. Also, Also, a minimal a minimal number number of blade of blade leads leads to small to wettedsmall wetted area and area friction and loss, friction which loss, requiring which torequiring be considered to be considered when the fanwhen pressure the fan is pressure low. is low. InIn accordanceaccordance with the matching methods ①1 ,, thethe bladeblade numbernumber ofof thethe tiptip turbineturbine isis limitedlimited byby thethe fan.fan. Then, the turbine blade can merge with the fan bladeblade slightlyslightly andand maintainmaintain thethe structurestructure constraintconstraint whichwhich is is positive positive for for the the blade blade strength. strength. However, However, being being mounted mounted at the at tip the of tip the of fan the blade fan whereblade wherethe diameter the diameter is large, is the large, solidity the ofsolidity the tip of turbine the tip is low. turbine Preliminary is low. Preliminary calculation shows calculation that theshows solidity that ofthe the solidity tip turbine of the can tip beturbine at approximately can be at approximately 0.6–0.7 which 0.6 is– much0.7 which less thanis much a conventional less than a turbineconventional (>1.0) turbine [20]. This (>1.0) low [ solidity20]. This problem low solidity may problem complicate may the complicate formation the of aformation typical flow of a path typical for turbineflow path blades. for turbine This topic blades. will This be discussed topic will subsequently.be discussed subsequently.
5.5. Numerical VerificationVerificationof of thetheADFTT ADFTTRotor Rotorand and DiscussionsDiscussions Currently, CFD CFD plays plays an an important important role role in the in aerodynamic the aerodynamic design design of turbo of machines. turbo machines. To evaluate To theevaluate aerodynamic the aerodynamic design of the design ADFTT, of the an innovativeADFTT, an and innovative uncertain and section, uncertain the ADFTT section, rotor the (including ADFTT therotor stator (includin of theg tipthe turbine), stator of is the numerically tip turbine), certificated is numerically through certificated the CFD method. through Inthe addition, CFD method. for an exploratoryIn addition, design,for an exploratory the air-bleeding design, channel the air is a-bleeding new component channel andis a requiresnew component verification. and However, requires consideringverification. that However, the channel considering is a positive that pressure the channel gradient, is a positive and the flow pressur Mache gradient, number can and be the limited flow atMach a low number level, wecan estimate be limited that at the a low difficulty level, ofwe designing estimate thisthat channelthe difficulty is far lessof designing than the ADFTT this channel rotor, andis far we less will than verify the ADFTT its design rotor, in a and follow-up we will work. verify its design in a follow-up work.
5.1.5.1. CFD Method Numeca, the commercialcommercial CFDCFD software,software, includingincluding meshingmeshing modulemodule AutogridAutogrid andand computingcomputing modulemodule Fine, is usedused forfor numericalnumerical simulation.simulation. An adiabatic and no no-slip-slip wall condition is adopted, andand aa periodicperiodic boundaryboundary isis setset forfor bladeblade passage,passage, therebythereby ignoringignoring thethe clearanceclearance ofof aa rotorrotor blade.blade. The flow flow field field was was calculated calculated using using a a 3D 3D Reynolds Reynolds Navier-Stokes Navier-Stokes equations equations methodmethod withwith aa space-centralspace-central difference difference scheme scheme and four-step and four time-step advancement time advancement implemented implemented using the Runge-Kutta using the method.Runge-Kutta There method. are several There turbulence are several models turbulence such as Baldwin-Lomax models such model, as Baldwin Spalart-Almaras-Lomax modelmodel,, modelSpalart and-Almaras so on. model, Also, the model turbulence and so can on. be Also, fully the resolved turbulence using can direct be numericalfully resolved simulation using (DNS)direct ornumerical partially simulation resolvedusing (DNS) large or partially eddy simulation resolved using (LES). large However, eddy simulation the computationally (LES). However, expensive the ofcomputationally DNS and LES preventsexpensive them of DNS from and being LES used prevents as a design them from tool. b Theeing mesh used resolution as a design and tool. quality The requirements for this class of turbulence models are also less stringent than the ones required in DNS or
LES. So those turbulence models are still wildly used. In this paper, the turbulence model is standard Spalart–Allmaras model, which is a one-equation model for the modified turbulent kinematic viscosity Energies 2018, 11, x FOR PEER REVIEW 9 of 16
mesh resolution and quality requirements for this class of turbulence models are also less stringent than the ones required in DNS or LES. So those turbulence models are still wildly used. In this
Energiespaper, 2018the, 11turbulence, 3350 model is standard Spalart–Allmaras model, which is a one-equation model9 of 16 for the modified turbulent kinematic viscosity . As one-equation model, Spalart–Almaras model shows good agreement with airfoils [21] which are the basic of the blade of impellers and requires νeless. As computation one-equation resources model, Spalart–Almaras than other two model-equation shows models. good agreementThe formulation with airfoils of the [ 21 mod] whichel is arethe thetransport basic ofequation the blade: of impellers and requires less computation resources than other two-equation models. The formulation of the model is the transport equation: D1 2 P D [ (( ) cb2 ( ) ] , (12) Dν Dt 1 e = P − D + [∇·((ν + ν)∇ν + c (∇ν)2], (12) Dt σ e e b2 e where / is the laminar kinematic viscosity. The production and wall destruction terms whereread: ν = µ/ρ is the laminar kinematic viscosity. The production and wall destruction terms read: P c(1 f ) S P = cb1(bt112− ft2)Seνe,, (13)(13)
cc ν 2 = ( − bb11 )[ e] 2 D Dcw(1 cwf12w f wf ft t2 )[ ] .. (14)(14) κ22 d More details of such turbulent model could be obtained from the reference [21]. More details of such turbulent model could be obtained from the reference [21]. The Spalart-Allmaras model assumes that the mesh is sufficiently refined close to the wall The Spalart-Allmaras model assumes that the mesh is sufficiently refined close to the wall surfaces with the non-dimensional wall distance y+~1. However, in engineering application, it is surfaces with the non-dimensional wall distance y+~1. However, in engineering application, it is difficult to insure all the mesh close to the wall surfaces with y+~1. So the use of y+ within 1–10 is a difficult to insure all the mesh close to the wall surfaces with y+~1. So the use of y+ within 1–10 is a compromise and suggested in the Numeca manual. In our further work, the experiment study will compromise and suggested in the Numeca manual. In our further work, the experiment study will be adopted, so the deviations between calculation performance (with this simulation setting) and be adopted, so the deviations between calculation performance (with this simulation setting) and actual performance could be known. Reference [22] checks the reliability of the calculation method actual performance could be known. Reference [22] checks the reliability of the calculation method and software, in which the fan has the similar rotating speed and size with the fan in this research. and software, in which the fan has the similar rotating speed and size with the fan in this research. The boundary conditions are set as follows: the total pressure of the inlet of tip turbine is The boundary conditions are set as follows: the total pressure of the inlet of tip turbine is 108,624 Pa, 108,624 Pa, the total temperature is 440 K, and static pressure of its outlet is 101,300 Pa; the total the total temperature is 440 K, and static pressure of its outlet is 101,300 Pa; the total pressure of the pressure of the inlet of air-driven fan is 101,325 Pa, the total temperature is 293 K, and static pressure inlet of air-driven fan is 101,325 Pa, the total temperature is 293 K, and static pressure of its outlet is of its outlet is 101,300 Pa; the adiabatic and non-sliding solid wall treatment method is adopted; the 101,300 Pa; the adiabatic and non-sliding solid wall treatment method is adopted; the blade surface blade surface and hub surface are set as rotating boundary, the casing is static boundary, and the blade and hub surface are set as rotating boundary, the casing is static boundary, and the blade passage is passage is set as periodic boundary. set as periodic boundary.
Software: Numeca Grid Number of Tip-Turbine: Approximately 1,130,000 Mathematical model: Turbulent Navie-Stokes
Modelling of turbulence: Spalart-Alimaras
Wall y+ ≈ 1~ 10
path ow e fl bin Tur ength: mm Chord L 955 m r= ≈310m
mm ath 70 w p r=8 flo Fan
m Grid Number of Fan: 60m r=4 Approximately 1,300,000
Figure 4.4. MediumMedium Grid of the ADFTT Rotor (including the stator of the tip turbine).turbine).
To show show the the grid grid independence, independence, three three number number grids (i.e., grids coarse, (i.e., medium coarse, and medium fine) are and constructed fine) are forconstructed the fan and for tip the turbine. fan and The tip numbers turbine. of The cells numbers for the fan of is cells 670,000 for the (coarse fan grid), is 670,000 1,300,000 (coarse (medium grid), grid) and 2,230,000 (fine grid) respectively. The numbers of cells for the tip turbine is 740,000 (coarse grid), 1,130,000 (medium grid) and 1,560,000 (fine grid) respectively. Figure4 demonstrates the merged grid of the fan and tip turbine (medium grid). Figure5 exhibits the results of grid number Energies 2018, 11, x FOR PEER REVIEW 10 of 16
Energies1,300,0002018 (medium, 11, 3350 grid) and 2,230,000 (fine grid) respectively. The numbers of cells for the10 of tip 16 turbine is 740,000 (coarse grid), 1,130,000 (medium grid) and 1,560,000 (fine grid) respectively. Figure 4 demonstrates the merged grid of the fan and tip turbine (medium grid). Figure 5 exhibits independencethe results of grid checking. number The independence mass flow rates checking. of both The components mass flow remain rates theof both same components with the increase remain in the numbersame with of grids,the increase whereas in the the efficiencies number of of grids, both componentswhereas the are efficiencies nearly unchanged. of both components The inlet Mach are numbernearly unchanged. and pressure The ratio inlet are Mach also number approximately and pressure the same ratio with are different also approximately grids for both the components. same with Thedifferent differences grids for in theboth calculation components results. The are differences nearly neglected. in the calculation For the rapid results preliminary are nearly design, neglected. less gridFor the numbers rapid prelimina will savery plenty design of, timeless grid on the numbers computation. will save However, plenty of a time suitable on the number computation. of grids willHowever, help reduce a suitable the deviationnumber of caused grids will by the help difference reduce the in thedeviation post processing caused by of the CFD difference results, soin the mediumpost processing grid is chosenof CFD forresults, the present so the medium simulation. grid is chosen for the present simulation.
Efficiency (%) Massflow rate (kg/s) Inlet Mach Pressure ratio
Coarse Medium Fine Coarse Medium Fine
(a) (b)
Figure 5.5. Grid independence results results.. ( (aa)) Efficiency Efficiency and and mass mass flow flow rate; rate; ( (b)) Inlet Mach number and pressure ratio.
5.2. CFD Results and Discussions
5.2.1. Tip Turbine 5.2.1. Tip Turbine The blade number of stator is 29 and the blade number of rotor is 12 (equals to the fan rotor). The blade number of stator is 29 and the blade number of rotor is 12 (equals to the fan rotor). Given the limitation in the blade number from an air-driven fan, the solidity of the tip turbine is too Given the limitation in the blade number from an air-driven fan, the solidity of the tip turbine is too low (approximately 0.6–0.7) and cannot establish an integrated flow path by its adjacent blades if the low (approximately 0.6–0.7) and cannot establish an integrated flow path by its adjacent blades if the turbine is designed through a normal design process of turbomachinery. Figure6a displays the Mach turbine is designed through a normal design process of turbomachinery. Figure 6a displays the number contour of the tip turbine which is designed through the normal process, and its isentropic Mach number contour of the tip turbine which is designed through the normal process, and its efficiency is only approximately 70%. The stator functions well and accelerates the flow to a high isentropic efficiency is only approximately 70%. The stator functions well and accelerates the flow to subsonic speed, but the rotor does not work well. The throat of the rotor flow path seems to be formed a high subsonic speed, but the rotor does not work well. The throat of the rotor flow path seems to be at Position A which is outside the flow path. The flow field shows a serious flow separation at the formed at Position A which is outside the flow path. The flow field shows a serious flow separation suction side of the blade and a large supersonic area at the downstream of the blade. Consequently, at the suction side of the blade and a large supersonic area at the downstream of the blade. the conventional design method is an unsuitable solution for the tip turbine. The main reason is that Consequently, the conventional design method is an unsuitable solution for the tip turbine. The blade solidity is too low to constrain airflow. In other words, the airflow deflects a large angle in this main reason is that blade solidity is too low to constrain airflow. In other words, the airflow deflects low solidity turbine and separation occurs in the latter part of the suction surface. a large angle in this low solidity turbine and separation occurs in the latter part of the suction surface.
Energies 2018, 11, 3350 11 of 16 Energies 2018, 11, x FOR PEER REVIEW 11 of 16
1.6 1.6
1.4 1.4
1.2 1.2
A 1.0 A 1.0
0.8 0.8
0.6 0.6 B 0.4 0.4
0.2 0.2 0 0 (a) (b)
FigureFigure 6. 6Mach. Mach Number Number Contour Contour of of the the tiptipturbine turbine (50%(50% span).span). (a) Conventional design; design; ( (bb)) H Highigh reactionreaction design. design.
TheThe reactionreaction turbine turbine is iswildly wildly used used in gas in turbine gas turbine area for area its positive for its positive pressure pressuregradient in gradient rotor [23 in]. rotorHigh [ 23 reaction]. High reaction turbine turbinehas two has benefits two benefits:: one is one its is positive its positive pressure pressure gradient; gradient; the the other other is isthe the reducingreducing the the deflectiondeflection angleangle ofof flowflow inin thethe rotor. Both of the the two two benefits benefits will will make make it it dif difficultficult for for airflowairflow to to separate. separate. However, However, in high-reaction in high-reaction turbine, turbine, the relative the relative velocity velocity in rotor in will rotor be increased, will be whichincreased, also limits which the also improvement limits the of improvement rotor’s efficiency, of rotor’s even cause efficiency, transonic even or supersonic cause transonic problems. or Sosupersonic high-reaction problems. turbine So is high seldom-reaction applied turbine in practice. is seldom applied in practice. ToTo relieve relieve the the problem problem caused caused by by such such a a low low solidity of the turbine, a a high high reaction reaction design design methodmethod isis adoptedadopted inin tiptip turbine.turbine. The relative flow flow angle angle between between the the air air and and the the axial axial direction direction in in thethe rotor rotor increases, increases, andand thethe turningturning angle o off the air in the rotor can can be be controlled controlled under under a alow low level. level. TheThe blade blade profile profile isis improved improved throughthrough aa directdirect problemproblem method. The The blade blade profile profile near near the the trailing trailing edgeedge must must be be optimisedoptimised toto ensureensure thatthat thethe streamline at Position A A is is nearly nearly parallel parallel with with the the pr profileofile ofof a a suction suction sideside atat PositionPosition B.B. FigureFigure6 b6b presents presents thethe MachMach numbernumber contourcontour of the tip turbine that usesuses the the aero-extending aero-extending bladeblade design.design. The physical throat seems seems to to be be located located at at Position Position A, A, but but the the flowflow field field shows shows that that the the aero aero throat throat is is located located at at Position Position B. B. Thus, Thus, the the blade blade solidity solidity is increasedis increased if theif the length of the blade is calculated until at least Position B with so called ‘aero-extending method’. length of the blade is calculated until at least Position B with so called ‘aero-extending method’. The ‘aero-extending’ seems to be an apparent phenomenon of high-reaction design with The ‘aero-extending’ seems to be an apparent phenomenon of high-reaction design with trailing trailing edge profile optimization. In addition, to reduce the impact of relative speed increase, the edge profile optimization. In addition, to reduce the impact of relative speed increase, the height of height of the turbine flow path is increased in the high reaction method (Figure 4 is the final grid of the turbine flow path is increased in the high reaction method (Figure4 is the final grid of this high this high reaction design). The increasing of path height will brings two benefits, one is decreasing reaction design). The increasing of path height will brings two benefits, one is decreasing the flow the flow velocity in stator and rotor (for the axial velocity is decreased), the other is reduce the flow velocity in stator and rotor (for the axial velocity is decreased), the other is reduce the flow velocity velocity of the outlet of stator and let the air accelerate more in rotor. What’s more, the tip turbine is of the outlet of stator and let the air accelerate more in rotor. What’s more, the tip turbine is not a not a high load or high pressure ratio drop turbine, so the relative velocity in rotor could be high load or high pressure ratio drop turbine, so the relative velocity in rotor could be controlled at an controlled at an acceptable level when high reaction turbine design is adopt. Though more losses acceptable level when high reaction turbine design is adopt. Though more losses will be produced will be produced due to the increasing of relative velocity of airflow, the improvement of turbine due to the increasing of relative velocity of airflow, the improvement of turbine performance is still performance is still effective. The working conditions of the ADFTT make the low solidity tip effective. The working conditions of the ADFTT make the low solidity tip turbine to be a special turbine to be a special problem and could be fixed by this method. problem and could be fixed by this method.
Figure7 shows more details of the flow field. The leading edge of hub is set as 0% section and the trailing edge of hub is set as 100% section. The 20% section is almost belongs to high pressure region 1 , and most area of 100% section belongs to low pressure region 2 . The high pressure region only covers the pressure side of blade, and the low pressure region only covers the suction side of the next blade. The air in the flow path of the rotor traverses two periods: Firstly, high pressure region, the flow is accelerated before Position A. Secondly, low pressure region, the flow is continue accelerated in the area between Positions A and B (the same with Figure6b). In conventional turbine, the air usually flows through high pressure region (near suction side) or low pressure region (near suction side). However, in this tip turbine, the air flows through the high pressure region into the low pressure
Energies 2018, 11, 3350 12 of 16 region. The flow field seems regular with only two small separations. One (S1) locates at the corner of tip near the leading edge. This separation is mainly caused by flow around leading edge (yellow stream lines). The other (S2) locates at the corner of hub near the trailing edge. This separation is mainly caused by the convolution of low velocity fluid of end wall. These two separations may be improvedEnergies 2018 in, the11, x further, FOR PEER so REVIEW that the efficiency of tip turbine could be further enhanced. 12 of 16
20% Section STATOR
n io t 60% Section c S1 e ir d 100% Section w lo F
150% Section
S2 2 1
ROTOR
0 20,000 40,000 60,000 80,000 100,000 120,000 140,000 160,000
FigureFigure 7. 7.Static Static pressure pressure contour contour of of the the cross cross stream stream of of tip tip turbine turbine at at different different sections. sections.
FigureFigure8 shows 7 shows the more blade details loadings of ofthe tip flow turbine field at. differentThe leading spans. edge The of blade hub is loadings set as 0% is almost section the and samethe withtrailing little edge differences of hub is at set leading as 100% edge section. and position The 20% B. The section pressure is almost on suction belongs side to clearly high pressure shows theregion accelerate ①, and area most from area leading of 100% edge section (near position belongs A) to to low the pressure middle of region the blade ②. (near The high position pressure B). Thisregion phenomenal only covers meets the thepressure flow field side mentionedof blade, and above. the low After pressure position region B, the o airnly meets covers the the adverse suction pressuresideEnergies of gradient 2018 the , next11, xwhich FOR blade. PEER could The REVIEW beair one in theof the flow reasons path for of the corner rotor traverses separation two at the periods trailing: Firstly, edge. 13high of 16 pressure region, the flow is accelerated before Position A. Secondly, low pressure region, the flow is continue accelerated in the area between Positions A and B (the same with Figure 6b). In conventional turbine, the airSta usuallytic pressur eflows (Pa) through high pressure region (near suction side) or low
pressure region (near suction160,00 0 side). However, in this tip turbine, the air flows through the high pressure region into the low pressure region. The flow field seems regular with only two small separations. One (S1) locates140, 0at00 the corner of tip near the leading edge. This separation is mainly caused by flow around leading120,000 edge (yellow stream lines). The other (S2) locates at the corner of hub near the trailing edge. This separation is mainly caused by the convolution of low velocity fluid 100,000 of end wall. These two separations may be improved in the further, so that the efficiency of tip turbine could be further enhanced80,000 . e ers v re Figure 8 shows the blade60,000 loadings of tip turbine In at differentsu spans. The blade loadings is res nt almost the same with little differences at leading edge andp positiondie B. The pressure on suction side 40,000 accelerate gra clearly shows the accelerate area from leading edge (near position A) to the middle of the blade A B (near position B). This phenomenal20,000 meets the flow field mentioned above. After position B, the air meets the adverse pressure gradient which could be one of the reasons for the corner separation at Leading edge Trialing edge the trailing edge. BesideFigureFigures, as 8. aBlade 8 result. Blade loading ofloading the of hightip of tip turbine reaction turbine at different atdesign different, spans the spans absolute (0% (0% for hubfor angle hub and andof 100% the 100% forflow shroud).for atshroud) outlet. is 14.7° (angle between the airflow direction and the axial direction, the same below). The normally ◦ designed5.2.Besides,2. Air turbine-Driving as a result could Fan of be the 0 high°. The reaction flow direction design, theangle absolute must be angle turned ofthe in the flow additional at outlet isstator 14.7 to (anglerestore between some the thrust airflow. directionNevertheless and, the given axial direction, the considerable the same below). efficiency The normally improvement, designed the The blade ◦ number of fan rotor is 12 (equal to the turbine rotor). Since the pressure ratio of turbine‘aero-extending could be 0’ .tip The turbine flow direction blade remains angle must impressive be turned as in a the new additional method of stator turbine to restore design. some The air-driven fan is low, more blades will lead to more friction loss and weight, but will not effectively thrust.perform Nevertheless,ance of this given tip turbine the considerable is summarised efficiency in Table improvement, 2. The isentropic the ‘aero-extending’ efficiency of this tip turbine turbine is improve the fan efficiency. Figure 9a illustrates the detailed flow field of the air-driven fan. There approximately 85.9% and satisfies the requirement of turbine performance. The design technique of are two small separations (red and purple stream lines) near the end wells. Similarly, these the ‘aero-extending’ tip turbine blade will be investigated carefully in future research. separations are both caused by the convolution of low velocity fluid of end wall. There is a small supersonic region at the tip of the leading edge, which is caused by the flow around leading edge. Figure 9b depicts the relative Mach number contour of the fan rotor at a 90% height span and shows the supersonic region. The relative Mach number of the entire flow field is nearly subsonic, and no flow separation or shock wave appears in the flow field. The main losses are friction and trailing wake. The relative flow direction is slightly turned by the fan rotor from 37.8° to 26.3°, and the absolute flow angles changes from 0° to 18°. This indicates that the circumferential velocity is minimal and the pressure ratio is low. In other circumstances, the air-driven fan rotor seems similar to a propeller because its solidity is much less than a conventional fan.
10% Section
50% Section
90% Section (a) (b)
Figure 9. Flow Field Details of the Air-Driven Fan Rotor. (a) Mach number contour of the cross stream at different sections; (b) Mach number contour of at a 90% span.
Figure 10 shows the blade loading of air-driven fan at different span. The blade is a fore-loading blade. The loading at the front area reduces from hub to shroud, in order to reduce the acceleration of the flow near the tip. This is the benefit of the sweep blade. Since the circumferential velocity of the blade is increased with the increase of radius, faster acceleration at tip could lead to
Energies 2018, 11, x FOR PEER REVIEW 13 of 16
Static pressure (Pa)
160,000
140,000
120,000
100,000
80,000
se er nv re Energies 2018, 11, 3350 60,000 I su 13 of 16 res nt p die 40,000 accelerate gra A B blade remains impressive2 as0,0 a00 new method of turbine design. The performance of this tip turbine is summarised in Table2. The isentropic efficiency of this turbine is approximately 85.9% and satisfies Leading edge Trialing edge the requirement of turbine performance. The design technique of the ‘aero-extending’ tip turbine blade will be investigatedFigure 8. Blade carefully loading in of future tip turbine research. at different spans (0% for hub and 100% for shroud). 5.2.2. Air-Driving Fan 5.2.2. Air-Driving Fan The blade number of fan rotor is 12 (equal to the turbine rotor). Since the pressure ratio of The blade number of fan rotor is 12 (equal to the turbine rotor). Since the pressure ratio of air-driven fan is low, more blades will lead to more friction loss and weight, but will not effectively air-driven fan is low, more blades will lead to more friction loss and weight, but will not effectively improve the fan efficiency. Figure9a illustrates the detailed flow field of the air-driven fan. There are improve the fan efficiency. Figure 9a illustrates the detailed flow field of the air-driven fan. There two small separations (red and purple stream lines) near the end wells. Similarly, these separations are are two small separations (red and purple stream lines) near the end wells. Similarly, these both caused by the convolution of low velocity fluid of end wall. There is a small supersonic region separations are both caused by the convolution of low velocity fluid of end wall. There is a small at the tip of the leading edge, which is caused by the flow around leading edge. Figure9b depicts supersonic region at the tip of the leading edge, which is caused by the flow around leading edge. the relative Mach number contour of the fan rotor at a 90% height span and shows the supersonic Figure 9b depicts the relative Mach number contour of the fan rotor at a 90% height span and shows region. The relative Mach number of the entire flow field is nearly subsonic, and no flow separation or the supersonic region. The relative Mach number of the entire flow field is nearly subsonic, and no shock wave appears in the flow field. The main losses are friction and trailing wake. The relative flow flow separation or shock wave appears in the flow field. The main losses are friction and trailing direction is slightly turned by the fan rotor from 37.8◦ to 26.3◦, and the absolute flow angles changes wake. The relative flow direction is slightly turned by the fan rotor from 37.8° to 26.3°, and the from 0◦ to 18◦. This indicates that the circumferential velocity is minimal and the pressure ratio is low. absolute flow angles changes from 0° to 18°. This indicates that the circumferential velocity is In other circumstances, the air-driven fan rotor seems similar to a propeller because its solidity is much minimal and the pressure ratio is low. In other circumstances, the air-driven fan rotor seems similar less than a conventional fan. to a propeller because its solidity is much less than a conventional fan.
10% Section
50% Section
90% Section (a) (b)
FigureFigure 9.9.Flow Flow Field Field Details Details of ofthe the Air-Driven Air-Driven Fan Fan Rotor. Rotor (a) Mach. (a) Mach number number contour contour of the cross of the stream cross atstream different at different sections; sections (b) Mach; (b number) Mach n contourumber c ofontour at a 90% of at span. a 90% span.
FigureFigure 10 10 shows shows the the blade blade loading loading of air-driven of air-driven fan at different fan at different span. The span. blade is The a fore-loading blade is a blade.fore-loading The loading blade. atThe the loading front area at the reduces front area from reduces hub to shroud,from hub in to order shroud, to reduce in order the to acceleration reduce the ofacceleration the flow near of the the flow tip. near This the is the tip. benefit This is of the the benefit sweep of blade. the sweep Since blade the circumferential. Since the circumferential velocity of thevelocity blade of is the increased blade is with increase the increased with the of radius,increase faster of radius acceleration, faster acceleration at tip could leadat tip to could higher lead flow to velocity and result in more losses, even shock wave. The profile of air-driven fan could continually be modified to get a more uniform loading and reduce the supersonic region near the leading edge.
Table 2. CFD Results of the air-driven fan and tip turbine.
Parameters Fan Rotor Tip Turbine Efficiency 89.7% 85.9% Pressure ratio (drop) 1.09 1.51 Mass flow rate (kg/s) 325 79.2 Power (kW) −2544 2654 Torque (N·m) −12,151 12,672 Energies 2018, 11, x FOR PEER REVIEW 14 of 16
higher flow velocity and result in more losses, even shock wave. The profile of air-driven fan could continually be modified to get a more uniform loading and reduce the supersonic region near the leading edge.
Table 2. CFD Results of the air-driven fan and tip turbine.
Parameters Fan Rotor Tip Turbine Efficiency 89.7% 85.9% Pressure ratio (drop) 1.09 1.51 Mass flow rate ( kg/s ) 325 79.2 Power ( kW ) −2544 2654 Energies 2018, 11, 3350 14 of 16 Torque ( N·m ) −12,151 12,672
Static pressure (Pa)
100,000
80,000
60,000
40,000
20,000
Leading edge Trialing edge
FigureFigure 10. 10Blade. Blade Loading Loading of Air-drivenof Air-driven Fan Fan at Differentat Different Span Span (0% (0% for for hub hub and and 100% 100% for for shroud). shroud). 5.3. Comparison between the Prototype Turbofan and the Demo ADFTT 5.3. Comparison between the Prototype Turbofan and the Demo ADFTT With the CFD results of the ADFTT, the prototype turbofan and the preliminary ADFTT can then With the CFD results of the ADFTT, the prototype turbofan and the preliminary ADFTT can be compared to indicate the performance advantages of the ADFTT. In Table3, when the ADFTT is then be compared to indicate the performance advantages of the ADFTT. In Table 3, when the applied to the prototype turbofan, the BPR can be improved by 413%, the thrust can be improved by ADFTT is applied to the prototype turbofan, the BPR can be improved by 413%, the thrust can be 84.0% and the SFC can be reduced by 45.6%. However, these results do not denote that the performance improved by 84.0% and the SFC can be reduced by 45.6%. However, these results do not denote that or the efficiency of the ADFTT is superior to the GTF or the conventional turbofan at the same BPR. the performance or the efficiency of the ADFTT is superior to the GTF or the conventional turbofan In fact, in accordance with Equation (7), the equivalent efficiency of the ADFTT can be calculated. at the same BPR. In fact, in accordance with Equation (7), the equivalent efficiency of the ADFTT can Therefore, the ADFTT, as the BPR ‘amplifier’, can truly significantly improve the performance of the be calculated. Therefore, the ADFTT, as the BPR ‘amplifier’, can truly significantly improve the prototype turbofan. performance of the prototype turbofan. Table 3. Comparison between the Prototype Turbofan and the Preliminary ADFTT. Table 3. Comparison between the Prototype Turbofan and the Preliminary ADFTT. Parameters Prototype Turbofan Preliminary ADFTT Alteration (%) Parameters Prototype Turbofan Preliminary ADFTT Alteration (%) BPRBPR 5.35.3 27.227.2 +413.0 +413.0 ThrustThrust (kN) ( kN ) 25.525.5 48.548.5 +84.0 +84.0
AfterAfter summarising summarising the thermodynamic the thermodynamic analysis, analysis, aerodynamic aerodynamic design and design numerical and simulation, numerical thesimulation, ADFTT is concluded the ADFTT as is a feasible concluded means as for a feasible improving means the performancefor improving of the a turbofan. performance Further of a workturbofan. must still Further beconducted work must to still develop be conducted the detailed to develop aerodynamic the detailed designs aerodynamic of components designs and of multidisciplinarycomponents and design multidisciplinary methods of the design ADFTT. methods of the ADFTT.
6. Conclusions6. Conclusions This research presents a conceptual design of an ADFTT for a high-bypass propulsion system. Through the work discussed above, the following conclusions can be drawn:
(1) A concept of a turbofan with an ADFTT is presented. This machine can solve the incompatibility problem of the fan and LTP, and is a promising device for high-bypass propulsion systems, especially for cost-controllable and small/mid-size aircraft. (2) Theoretical thermodynamic analysis of the ADFTT is proposed. The equivalent efficiency is also introduced to evaluate the energy transfer process. The ADFTT significantly improves the bypass of the prototype turbofan and increase the thrust. (3) An ‘aero-extending’ design can resolve the low solidity problem of the rotor of the tip turbine and ensure the positive performance. The aero-extending blade can extend the blade aerodynamically and achieve the design target of the efficiency. Energies 2018, 11, 3350 15 of 16
(4) This research aims to perform the exploratory design of the ADFTT, including the CFD results of the rotor of the ADFTT. The results of the demo work show the potential feasibility of the ADFTT, and further improvements can be expected. In comparison with the prototype turbofan, the performance of the CFD results shows that the ADFTT can improve the thrust by 84.0% and reduce the SFC by 45.6%.
Author Contributions: Conceptualisation, G.H. and X.X.; methodology, X.X.; validation, X.X.; formal analysis, G.H.; investigation, X.X., W.L. and L.L.; resources, G.H.; data curation, X.X.; writing—original draft preparation, X.X.; writing—review and editing, G.H. and C.X.; visualisation, X.X. and G.H.; supervision, G.H.; project administration, G.H.; funding acquisition, G.H. Funding: This research was funded by the National Natural Science Foundation of China (Grant number 51176072). Acknowledgments: The authors express their gratitude to the Jiangsu Province Key Laboratory of Aerospace Power System (affiliated to the College of Energy and Power Engineering, Nanjing University of Aeronautics and Astronautics) for the technical support. The authors are also grateful to the team members of the College of Power and Energy of Nanjing University of Aeronautics and Astronautics for their cooperation. Conflicts of Interest: The authors declare no conflict of interest.
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