Abstract
Research in this paper is concerned on bay cooling of the new type of turboprop engine. The considered engine integration substantially increases the temperature inside the front nacelle compartment. In order to achieve the optimal temperature conditions for engine parts inside the nacelle for all operating regimes and engine/nacelle regime after aircraft landing at heat soak-back, a new bay cooling system is proposed. Proposed cooling system consists of standard NACA inlets at the front of the nacelle, two additional groups of gills on rear part of front nacelle compartment, Zone1, and standard nacelle gaps (around exhausts), plus spinner gap. The sizing approach in the given multicriteria task is based on given temperature ratings of the engine accessories. Moreover, for structural reasons, metal skin and stiffeners in the front part of the nacelle should be maintained below the defined temperature limits. Using a 3D CFD model of the front nacelle compartment, this compartment is analysed, utilizing the software ANSYS. The benchmark testing of the considered turboprop engine of ATP type, performed by manufacturer, is used for defining the all necessary boundary conditions in this research.
1. Introduction
Among other crucial regimes for testing the efficiency of the turboprop engine cooling system is the idle engine condition. Low and high idle levels are considered [1], to distinguish between the ground minimum generator revolutions and the flight minimum power, respectively. For the turboprop cooling system capability, the decisive condition on the ground is the low idle level. With respect to the ground idle condition, the following focus areas are to be addressed: (i)The impact of reverse thrust on the turboprop cooling efficiency [2](ii)Flight idle thrust required for taxiing on the ground [3](iii)Noise level from aircraft ground operations [4](iv)Combustion efficiency and emissions [5](v)Aircraft deicing [6]
As regards the first issue, the reverse thrust effect on the cooling deficiency is not significant with respect to the twisting of propeller blades such that the blade angle is smaller at the tip compared to the root. Hence, a limited negative angle applied to the whole blade makes only the outer part of blade (near tip) with the negative angle required for the reverse thrust. The second issue brings about another problem whether the ground idle case analysis is sufficient for the turboprop cooling design or not. Typically, cowl design must be frozen before clearing the integration for initial flight test campaign. In [3], an ideal turboprop engine design including cooling system design, particularly effective in idle conditions, is carried out. The third and fourth issues are coupled because the combustion efficiency and resulting emissions also influence the noise generated by the turboprop. These issues are considered to tackle for environmentally friendly turboprop operation not only on the ground but also in other flight conditions. The aircraft taxiing characterized by the turbopropeller derotation is modelled and validated on measured data to obtain an optimum ground aircraft handling [7]. Beside the efficient combustion, low emissions, noise suppression, and minimum thrust to taxi, this ground aircraft handling should also guarantee the turboprop engine cooling.
This research paper presents a design of turboprop cooling system that is able to cool the engine during the ground aircraft handling until the engine is shut down. This design concerns a passive cooling system of the new ATP turboprop engine with an additional air discharge. The passive cooling system is designed such that the airflow inside the nacelle of the turboprop engine cools the most sensitive accessories inside the nacelle. Formerly designed gearbox cooler can be found in Steiner [8]. Earlier, Segal [9] proposed a system that guarantees efficient ventilation both on the ground, at static conditions via free convection at the front and the rear areas of the engine bay, and in flight via unidirectional flow from the front inlets to the rear outlets. The modelling of heat transfer in the engine bay of the small airplane I-23 has been performed based on the control volume approach for heat and fluid flow as well as for thermal radiation [10]. In [11], the heat transfer model of the engine bay is simplified into 1D model of the nacelle described by a lumped thermal model together with a feasible k-ɛ turbulence model. In this research paper, both models are applied to the passive cooling system design for the nacelle and engine accessory parts. The performance of the ventilation and cooling system is investigated in case of the engine nacelle of a helicopter under hover and forward flight conditions [12]. In contrast to works [9–12], this research paper considers the engine with an additional hot air stream inside nacelle to change substantially temperature distribution and air flow pattern in the nacelle. Moreover, this research paper significantly extends the results of previous authors’ work [13], to the ground idle condition in order to minimize the time to shutting down the engine and to reduce environment exposure. The main goal of this research is to design a passive cooling system of the new ATP turboprop engine.
This research paper proposes novel cooling system of the new ATP turboprop engine as a result of collaborative research. As regards the engine type and size, heat sources, materials, and structures of which is made, this engine is similar to some extent to that engine from [10, 11]. The result is compared with the passive cooling system described in [11]. This paper is organized on the following way. Simplified 3D model of considered turboprop engine is given in Section 2. The next section includes the necessary boundary conditions and the most important engine parameters. In Section 4, the simulation results and their analyses are given in details. In Section 5, comparative CFD model study is presented, and finally, in the last section, the paper is concluded.
2. 3D Model of the Hot Part of the Turboprop Engine
Parts of the turboprop engine, which are located inside the engine cowl, operate in different temperature range. All of these elements in combination with engine and nacelle create a region of very high temperature that requires an appropriate cooling to provide a temperature within operating limits of all installed devices under the cowl. The problem of maintaining the required temperature field in hot engine compartment is more complex in the case of the new ATP engine, due to the existing additional heat sources.
In general, the two characteristic operating regimes after landing are considered. Immediately after aircraft landing, taxiing, and second regime when the engine keeps running, the propeller is feathered, thus providing minimum axial airstream past nacelle. In this phase, overheating of the front compartment will be prevented. At the very beginning, the used 3D model for both cases will be introduced.
The hot part of the turboprop engine consists of turbine section, combustion chamber, and gearbox in the front part of the engine compartment. Complexity of the interior nacelle geometry and model boundary are chosen, in order to achieve the most realistic results and satisfactory computation accuracy. The model boundaries are set on outside engine cowl surfaces and on outside engine surface under the cowl. The main goal in this research is to design the cooling system that will provide appropriate ventilation rate inside the nacelle that will ensure a proper function of all devices installed.
A 3D model of the bay cooling is shown in Figure 1. In the front part of the nacelle, the two NACA (National Advisory Committee for Aeronautics) inlets are proposed together with two pairs of the gills on the rear part of the nacelle. Each pair of inlet/outlet openings are placed symmetrically with respect to the turboprop engine as shown in Figure 1.
The dimension and placement of NACA inlets and spinner gap are taken from the given FTB (flying test bed) aircraft; hence, they cannot be changed in this investigation. The outlet ribs are designed for this new passive cooling system, and their dimensions are verified by computation on the case study. The model also includes exhaust gap between the nacelle and exhaust stacks. This gap has significant influence on the air flow inside the front nacelle part, in conjunction with spinner gap.
The Figure 2 shows the 3D mesh of considered bay cooling model. This mesh is consisting of approximately 1.4 million cells. The mesh quality is specified by using the standard mesh parameters and given as follows. (i)The range of the aspect ratio is 1.05 to 1.6(ii)Orthogonal grid quality is between 0.8 and 1 with an average value of 0.9(iii)Skewness is in the range from 0 to 0.5 with an average value of 0.25
The mesh size and quality introduced above are obtained after several iterations in the grid independent study. This study is performed in order to achieve a trade-off between accuracy and needed calculation time that is done on three different 3D meshes from 1 million up to 2.2 million cells, and the middle one is shown in Figure 2.
The results of the grid independent study for several types of the grid are given in Figures 3 and 4. In these figures, the results for the two extreme cases (coarse and finest grid) are given. The results for both specified meshes show similar calculation accuracy. The medium mesh with approximately 1.4 million cells is chosen. Calculation step size is set to 0.005 s, and this step ensures the convergence of calculation. Longer time step was used in later calculation phase, and minimum time step that provides the calculation stability is 0.1 s. The time step is limited due to the very high turbulence inside the cowl caused by additional PT jets.
3. Boundary Conditions
At the very beginning, before defining the boundary conditions, the main goals of this research should be repeated and emphasized. Three critical engine limitations (triple red point) that represent the most extreme working conditions will be referenced for defining the boundary conditions. The case study describes the situation after landing the aircraft, e.g., during the taxiing with given unfeathered propeller speed. Due to the natural engine heat sources, a sufficient passive cooling should be provided to ensure proper working condition for all accessories inside the nacelle. The benchmark testing of the considered turboprop engine of ATP type provided all the data necessary for boundary conditions in this research.
Experimentally measured data are used for defining the boundary surface temperature of gearbox and the hot part of engine including the combustion chamber. This surface temperature has a uniform profile along the hot part of the engine with small drop between combustion chamber and turbine part. Surrounding temperature of nacelle is set to 30°C. It is the highest average temperature in the considered FTB operating region.
In order to proceed towards representative boundary conditions in the given computational domain, axial and radial velocity profiles external to the nacelle were obtained from virtual blade model (VBM). The first set of assessed cases used plain approximate inlet velocity 15 m/s at the two NACA inlets and technical gaps behind the propeller and around the exhaust. This can be considered enough at concept-type design consideration.
4. Used Mathematical Models and Simulation Results
4.1. Case of Propeller Feathered
Due to the very high velocity rate regions with very turbulent character of the air flow inside the nacelle, the standard k-ε model is used. In these simulations, the classic heat convection model and simple Rosseland model, see ANSYS [14], as radiation model is used. As a solution method, the pressure velocity coupling with simple scheme is used in ANSYS environment. The second-order upwind method is used for the spatial discretization of all considered models. Initial simulation time step is 0.005 s with fixed stepping method. The initial temperature field is set as constant at 250°C.
As it can be seen from Figure 5, the axial air flow stream between additional hot inlets and rear ribs is provided. Since direct impingement of any airstream to engine surface shall be avoided, in this moment, it is necessary to check the structure of the internal flow thoroughly. Figure 6 supports that task displays the velocity distribution inside the hot engine part in two planes. Moreover, contours of the velocity magnitude show that the cold air stream from NACA inlets and spinner gap behind the propeller hub reaches the rear ribs and isolates the nacelle from the internal hot air. That enables to meet own nacelle structural requirements.
In Figure 7, the design feature that the substantial part of the heat from the additional jet is outflown through the outlet pairs of gills on the both sides of nacelle is presented. Analysing the temperature distribution at the rear part of this compartment the considerably higher temperature distribution is apparent as expected. Additionally, as shown in Figure 8, this distribution is evolved into hot vortex on the outflow.
Figure 9 shows that in the rear part of the nacelle, the temperature is considerably higher than in the front part, and this zone is cooled by two pairs of ribs on each side of the nacelle as mentioned in the previous section.
As also shown in Figure 8, the front part of the nacelle has a substantial lower temperature due to the cold air supplied by NACA inlets and ring gaps (spinner and stacks) in the front part. This is very important, since propeller system components are typically mounted in this area and shall be provided with uniform thermal loading. One can conclude that the very intensive surface stream close to the inner cowl surface decreases the surface temperature (see Figure 10) and keeps the desired temperature.
4.2. Case of Propeller Unfeathered
The second case, with engine running with propeller unfeathered after aircraft landing, is also analysed. The boundary conditions from the previous study are changed on NACA inlet and spinner gap. The input velocities on inlets are 20 m/s. Simulation of the second case gives the very similar results as in case of previous simulations. Thus, the temperature and velocity distributions are very close to the temperature and velocity distributions in the first case. In Figure 11, the velocity distribution in the second case is shown when the inlet velocity is 20 m/s at the NACA inlet and spinner gap.
5. Comparative CFD Model Study
To prove that the proposed cooling system for the new ATP type of turboprop engine is viable and efficient, a comparative study on CFD models is carried out. The CFD model from previous section is compared with the CFD model of the hot engine part coming from [11] for the ground idle condition. Both CFD models are considered for the case of propeller feathered. In Figure 12, the temperature distribution comparison along the hot engine part is presented.
Inspecting the comparison results in Figure 12, the cooling efficiency of the proposed cooling system results in average higher by 20% percent than this is in case of the comparative cooling system coming from [11]. In peak value, at length position 0.55 m, the cooling efficiency results higher even by 40% opposite to the 1D thermal model from [11]. With respect to the relative error between 1D and 3D thermal models approxl 10% in maximum (see [11]), the maximal cooling efficiency results higher by approxl 30%. The comparison recorded in Figure 12 shows the benefit of the proposed cooling system over the comparative passive cooling system due to that from [11]. Despite higher heat intensity (i.e., an additional hot air stream inside nacelle) of the proposed engine model, the cooling efficiency results in this model are higher than in case of the comparative model, except for the rear part of engine at length position greater than 0.6 m. As highlighted in the previous sections, the engine components located in the front part close to the propeller are the most sensitive to higher temperature. Generalization and practical impact of this comparative study are pointed out in the conclusions.
6. Conclusions
In this research paper, three case studies are performed. The first two case studies focused on CFD modelling, the new ATP-type engine running with propeller feathered and unfeathered after aircraft landing gives very similar results on temperature and velocity distribution inside the nacelle. These two studies verify the effect of the proposed passive cooling system, and the bay cooling model for the new ATP type of turboprop engine is achieved. The third case study is the comparative CFD model study proving the efficiency of the proposed cooling system higher enough than the cooling efficiency in the case of the comparative CFD model.
The main goal of this research is achieved because the axial air flow stream along the nacelle ensures the desired temperature distribution inside given tight twin-engine nacelle. This temperature distribution guarantees proper function of the installed undercowl airframe devices and engine accessories after aircraft landing. Primarily, the temperatures in the front part of the nacelle, Zone1, are maintained in the required limits. Compressor screen plenum, Zone2, and accessory gearbox compartment, Zone3, have been subject to separate design efforts.
The comparative CFD model study highlights practically superior cooling efficiency in case of the new ATP type of turboprop engine over other cases of engines similar to that in [10, 11], particularly when the cooling efficiency of the proposed cooling system is higher by 30% in peak value and 20% in average opposite to that efficiency of the cooling system from [11]. Benefit coming from this practical superiority is that the engine parts inside nacelle, especially those sensitive to higher temperature, will have longer faultless period, in general.
As it can be seen from the case studies and their simulation results, the proposed cooling and ventilation system is able to keep the temperature field inside predefined limits. Moreover, the presented study enabled to distribute dedicated instrumentation in areas of interest. Finally, in the paper, the proposed cooling system and its concept are proved to be efficient not only for the new ATP type of turboprop engine but also for similar ones.
The future activities will be concerned to investigate the temperature fields in flight regimes, using both CFD and airborne defensive aids suites (DAS).
Data Availability
This study contains results of an applied research. All the data used to support the findings of this study may only be provided with restrictions, i.e. only as anonymized data, due to commercial confidentiality.
Conflicts of Interest
The authors declare that they have no conflicts of interest.
Acknowledgments
The authors acknowledge the support from the ESIF, EU Operational Programme Research, Development and Education, Center of Advanced Aerospace Technology (CZ.02.1.01/0.0/0.0/16_019/0000826), and Faculty of Mechanical Engineering, Czech Technical University in Prague.