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Pagination not final (Please cite DOI) 1

ARTICLE A hybrid propulsion system for a high-endurance UAV: configuration selection, aerodynamic study, and gas turbine bench tests R. Capata, L. Marino, and E. Sciubba

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Abstract: In recent years, renewed interest in the development of unmanned aerial vehicles (UAVs) has led to a wide range of interesting applications in reconnaissance and surveillance. In these missions, the noise produced by propeller-driven UAVs is a major drawback, which can be partially solved by installing an electric motor to drive the propeller. While the evolution of high performance brushless motors makes electric propulsion particularly appealing, at least for small and medium UAVs, all electric propulsion systems developed to date are penalized by the limited range and endurance that can be provided by a reasonably sized battery pack. In this paper we propose a hybrid propulsion system based on a recently developed ultramicro gas–turbine (UMGT), which can be used to power an electric generator, providing a significant range and (or) mission time extension. The UMGT is undergoing operational testing in our laboratory, to identify the most suitable configuration and to improve performance: a new compact regenerative combustion chamber was developed and several tests are being carried out to reduce its weight and size so as to increase, all other things being equal, the vehicle payload. This paper aims to propose a high endurance UAV, by a preliminary configuration selection and aerodynamic study of its performance. Key words: UAV, ultramicro gas turbine, long-range drones, payload.

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Résumé : Au cours des dernières années, un regain d’intérêt dans l’élaboration de véhicules aériens sans pilote (UAV) a mené à une foule d’applications intéressantes en matière de reconnaissance et de surveillance. Or, le bruit produit par les UAV à hélice constitue un inconvénient majeur; ce bruit peut être atténué en partie en installant un moteur électrique actionnant l’hélice. Bien que l’évolution des moteurs sans balais à haut rendement rende la propulsion électrique particulièrement intéressante, du moins pour les UAV petits et moyens, tous les systèmes à propulsion électrique élaborés à ce jour sont défavorisés à cause de la faible portée/endurance qu’offre un bloc-batterie de taille raisonnable. Dans le présent article, nous proposons un système de propulsion hybride s’appuyant sur une ultra micro turbine à gaz (par la suite appelé UMGT) de conception récente, qui peut être utilisé pour alimenter un générateur d’électricité et procurer ainsi une prolongation appréciable du temps de vol lors d’une mission. À l’heure actuelle, l’UMGT fait l’objet d’essais opérationnels dans notre laboratoire afin d’identifier la configuration la plus appropriée et d’améliorer la performance: nous avons élaboré une nouvelle chambre de combustion régénératrice compacte et plusieurs essais sont en cours afin d’en réduire le poids et la taille de manière à augmenter, ceteris paribus, la charge utile du véhicule. L’objectif du présent article est de proposer un UAV à grande autonomie en effectuant une sélection de configuration préliminaire ainsi qu’une étude aérodynamique de sa performance. Mots-clés : UAV, ultra micro turbine à gaz, drones à longue portée, charge utile.

Received 3 April 2013. Accepted 19 December 2013. R. Capata, L. Marino, and E. Sciubba. Department of Mechanical and Aerospace Engineering, University of Roma Sapienza, Roma, Italy. Corresponding author: R. Capata, (e-mail: [email protected]). J. Unmanned Veh. Syst. 2: 1–20 (2014) dx.doi.org/10.1139/juvs-2013-0005

Published at www.nrcresearchpress.com/juvs on XX XX 2014.


J. Unmanned Veh. Syst. Vol. 2, 2014

1. Introduction

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The concept of remotely piloted or unmanned aerial vehicles (UAVs) was developed in the late 1970s, but it was only in the mid-1980s that the first UAVs completed their first official test flights. Since then, the technology has been substantially refined and major improvements in safety and reliability have been accomplished. In recent years, a renewed interest in the development of UAVs led to a wide range of interesting applications in the fields of reconnaissance and surveillance. The UAV roadmap 2000–2025 (U.S. Department of Defense 2001) highlights the importance of small and medium-sized devices for strategic missions. Strong concerns about the severity of global warming have increased the interest in so-called “green energy”, prompting worldwide efforts to reduce anthropogenic CO2 emissions. In line with such a global trend, electric-powered hybrid reconnaissance vehicles have already been commercialized and even electric-powered aircraft are under development (Logan et al. 2007; Kim et al. 2012; Schoemann and Hornung 2012). Today’s small UAVs have evolved both in their design and in the technology they make use of. The key technological issues involve power storage improvements, innovative motor design, avionics miniaturization, better liftto-drag ratio and stiffness characteristics. For electrically powered vehicles, the power storage system, usually a battery, represents by far the heaviest component. Improvements in power storage are therefore considered the best “window of opportunity” to decrease the overall weight of the vehicle and (or) improve the performance. A little reflection makes it immediately clear that there are practical limits to the use of secondary electric batteries as the UAV primary power source. For example, physiological limits on the part of the human “launcher” set a practical limit on the total vehicle weight for hand launching, and if the vehicle weight cannot grow, neither can its endurance: alternate energy storage systems, such as primary batteries, must be used. Another important design issue is the use of a propeller as propulsion device, which is a direct consequence both of the limits that many small airfields impose on the noise level, and of the need to minimize the acoustic signature. In addition to the design of a quiet and efficient propeller, the noise problem can be partially solved by installing an electric motor drive instead (Gur and Rosen 2009a, 2009b). The evolution of high-performance brushless motors makes electric propulsion particularly appealing, at least for small- and medium-sized UAVs, but all electric propulsion systems developed to date are characterized by a limited range and endurance. In spite of the perhaps over-enthusiastic statements made by some authors who claim complete sustainability and feasibility of small UAVs powered by small, hydrocarbon-fuelled internal combustion engines (ICE; Ronney 2005; Barnard 2006), some doubts remain about the practical implementation of such an option; some of the data in the preceding references appear at odds with each other and with the available data on commercial ICE. In fact, calculated or measured consumptions vary from 217 (Barnard 2006) to 395 (Himoinsa 2014) to 500 g/kWh (Teknel 2014), for an approximatly equal weight of device (500–600 N) for all the referenced sources. The choice of a hybrid propulsion system can overcome the endurance limitation and here we propose a solution based on a recently developed, ultramicro gas–turbine (UMGT) that can be used to power an electric generator (Capata and Sciubba 2006). A possible strategy to extend the range or endurance mission of the UAV consists of using the UMGT as the primary source of propulsive power for the cruise phase (the so-called “transfer to target”), while in the final approach, in which a quieter flight attitude is a demanding specification, the battery pack drives the propeller, and the UMGT is shut off. Other operational curves are also possible, and the calculations presented here are in fact based on a “minimum global weight profile” in which the UMGT is repeatedly turned on and off during the 12 h mission to recharge the battery. The main advantage of using a gas turbine (GT) group in lieu of a small ICE for electrogeneration consists in the substantially higher specific power of the former. If a “power density” is defined as the ratio between the weight of the device and its nameplate power, for a commercial GT group of about 2.5 kW class one obtains about 10 N/kW, while in the case of a hydrocarbon fuelled ICE of the same class (Ronney 2005; Capata and Sciubba 2006; Mitsubishi web site) this ratio is about 25 N/kW. Because the taxing constraint defined by the payload is a primary design specification for UAVs, the choice of a piston engine is far from rational. This consideration actually prompted the proposal discussed here of adopting a gas-turbine device as the thermal source. For example, considering only commercial models available on the market, the Jet-Cap prototype for model aircraft (about 5 kW) weighs about 50 N, versus a typical corresponding 5 kW ICE ranging between 100 and 150 N. Published by NRC Research Press

Pagination not final (Please cite DOI) Capata et al.

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Fig. 1. Geometry of the flying wing.

2. Aerodynamic and performance analysis

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This section describes the preliminary aerodynamic analysis of the design of our UAV, a necessary step for the correct sizing of the power plant. The aircraft conceptual design presented here is based on mission requirements that include a 12 h cruise phase at an average speed of 33.3 m/s at an operative altitude of 5000 m and for a dash speed of 45 m/s. The maximum take-off weight is 500 N, with a payload of 80 N. Under these assumptions, a flying wing configuration for the UAV was defined, with a length l of 2.0 m and a span b of 3.2 m. A system of elevons assures the pitch and roll motion while a double vertical tail, in which a pusher propeller is lodged, guarantees the yaw stability and control. In tailless configurations the adoption of elevons instead of the more traditional elevator— aileron system is usually preferred to increase the available surface for both of the different control functions. In fact almost all the spanwise space of the rear edge of the wing can be equipped with a movable surface that partially compensates for the lower leverage caused by the missing tail. Some details concerning the stability and control of such configurations can be found in Chao et al. (2010). The flying wing configuration has no fuselage drag, but it is known to be affected by a rather problematic trim and by an intrinsically low useful payload space (Capata et al. 2010). In the following sections we will show that, for the characteristics of the payload and of the propulsion system adopted here, these limitations can easily be overcome. The wing has a surface S = 4.0 m2, a root chord cr = 2.0 m, and a tip chord ct = 0.5 m. The taper ratio is λ = ct/cr = 0.25 and the aspect ratio A = b2/S = 2.56. The leading edge sweep angle, Λ, is 43.15°. A 3D sketch of the geometry of the wing is reported in Fig. 1. The twin tail consists of two identical trapezoidal surfaces with an area 0.2 m2 each, a span bt = 0.4 m, and root and tip chords equal to 0.8 and 0.2 m, respectively. The aspect ratio of the tail is 0.8, its taper ratio 0.25 and the leading edge sweep angle is 56°. Published by NRC Research Press



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The low mean value of the Reynolds number suggests the adoption of a low speed airfoil for the wing profile: here, the NASA/LANGLEY LS(1)-0417 (GA(W)-1) and NASA/LANGLEY LS(1)-0413 (GA(W)-1), both of which show good behaviour in terms of aerodynamic efficiency and pitching moments, were selected for the root and tip sections, respectively. A comparison against standard NACA 65 airfoil series is shown in McGhee and Beasly (1973), in which a more efficient behaviour in terms of lift and efficiency is clearly demonstrated. We approach the problem by considering first the nonviscous potential model to predict the lift characteristics and then complete the analysis with a simplified method to take into account the drag due to friction. This approach is quite standard in preliminary aerodynamic analyses (Roskam 1985), where a complete 3D viscous-flow simulation could be too expensive computationally, in the first design phase. The nonviscous aerodynamic characteristics of the UAV have been obtained by means of a vortex lattice method. We have chosen to follow this approach because of its proven effectiveness to analyse surfaces at low values of the angle of attack: it turns out that it also provides a good estimate of the lift and the induced drag of the aircraft. Details about the numerical procedure can be found in Bertin and Smith (1998) and Minotti and Sciubba (2010); here we shall only discuss the main points of the algorithm and show some of the most significant results. The code is based on a 3D solution of the potential flow around a complex body, with a model for a flexible wake and allows for the evaluation of the 3D forces acting on each panel, of the aerodynamic coefficients in both body and wing axis and of the stability derivatives with respect to angle of attack, angle of sideslip, angular rates, and rudder deflections. As an example, we report in Fig. 2 the vortex distribution over the wing and the twin tail. Different vortex distributions have also been studied and a good numerical convergence has been obtained with a distribution of 400 vortex panels on the wing and 100 panels on each vertical surface. In Fig. 2 the behaviour of the error on the evaluation of the lift coefficient for two different angles of attack (α = 1° and 10°) is shown as a function of the number, n, of the panels adopted on the wing. The error has been evaluated as 100 × ∣(CL(n)-CL(nmax))/CL(nmax)∣ It can be observed that the relative error falls below 1% when the number of panels is greater or equal to 400. For the specified cruising conditions, V = 33.3 m/s, h = 5000 m, Figs. 3 and 4 show the spanwise distribution of the local lift coefficient and the aerodynamic loading, respectively. Notice that the aerodynamic loading is smoothly distributed in the spanwise direction, thus minimizing the induced component of the aerodynamic drag. Figure 5 shows the distribution of the load coefficient, Δcp, on the wing. This quantity, defined as the local difference between the pressure coefficients on the upper and lower surfaces of the wing, has a discontinuity-free distribution and thus assures smooth behaviour in operation, particularly important at high angles of attack and in manoeuvring conditions. Figure 5 refers to two values of the angle of attack. The distributions are qualitatively similar with obvious differences in the absolute values corresponding to two very different flight conditions. Figure 6a corresponds to α = 1, while Fig. 6b has been obtained for α = 10: because of the high aerodynamic loading, the value of Δcp is much larger in this second case. The lift and drag characteristics can be derived from the geometrical and aerodynamic features of the aircraft by the polar equation CD = CD0 + k C2L , which provides the drag coefficient distribution as a function of the lift coefficient. For the UAV studied here, at the design operative conditions we have: ð1Þ

CD ¼ 0:02 þ 0:15C2L

We recall that, under level flight conditions, CL = W/(0.5ρV2S) and CD = D/(0.5ρV2S). The parasitic contribution CD0 to the drag coefficient in Eq. (1) has been calculated following the method of the equivalent flat-plate area as described in (Roskam 1985). In fact for a highly streamlined, aerodynamically clean shape the zero lift drag coefficient CD0 is mostly due to friction and only to a lesser degree to form drag. Following Hoerner (1965) and Paterson et al. (1973). This contribution can be written as CD0 ¼ Cf ðSwet =Sref ÞFF, where Cf is the friction factor, Swet the wetted surface, Sref the reference surface (plant form wing surface), and FF the form factor. For planar surfaces FF = 1 + 1.8(t/c) + 50 (t/c)4 with t/c the maximum thickness to chord ratio. The factor k, which affects the induced drag coefficient, CDi, is evaluated as k = 1/(πAe), where A is the aspect ratio and e is the Oswald factor assumed equal to 0.8, as suggested by Roskam (1985). To obtain reliable data on the drag coefficient the flow has been assumed to be turbulent over all of the UAV and Cf = 1.455/ [log(Rec)]2.58. Tables 1 and 2 summarize the main parameters, which drive the design of the propulsion system. In all calculations, a propeller efficiency of 0.7 was assumed. Published by NRC Research Press


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Fig. 2. Lift coefficient convergence versus the number, n, of panel adopted for the wing.

The mission requirements, following a general specification provided by a public call by the Italian Ministry of Defence, mandate that the payload be contained in a box of size 0.5 m × 0.5 m × 0.2 m and the range is imposed by the requirement of a data collection spanning an entire day. The payload was calculated approximately on the basis of the weight of the usual reconnaissance devices (cameras, scanners, etc.).

Published by NRC Research Press



Fig. 3. Discretization of the flying wing with the vortices distribution.

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Fig. 4. Lift coefficient of the wing sections versus spanwise distance.

The UAV total weight was imposed by the specification that it could be launched off the field by a portable launching sled transported and operated by only one person. The maximum thickness of the wing, at the longitudinal symmetry plane, is about 0.272 m. A check was conducted on the decreasing spanwise wing thickness, to ensure that there is enough space to accommodate the payload assembly. The remaining volume, that is, the internal volume of the flying wing minus the volume of the payload box, is then available for the propulsion plant, the fuel tanks, and the flight control and guidance system. Published by NRC Research Press


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Fig. 5. Aerodynamic loading versus spanwise distance.

3. The ultramicro turbo group

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3.1. Introduction This section describes the general engineering characteristics of the ultramicro turbo group proposed for the UAV. The compressor and turbine have been analysed and simulated in detail in previous papers (Capata and Sciubba 2009), and therefore only a brief description is provided here. The UMGT device described in the preceding references consists of a radial compressor, an air preheater, a combustion chamber (CC), and a radial turbine that together compose the thermal section; a double effect (reversible) electrical engine; a shaft and a set of high-speed bearings; and obviously all of the necessary auxiliaries like valves, controllers, and ducts, including the management unit. For safety reasons the cylindrical fuel tanks are external to the metallic UMGT enclosure. The feasibility of adopting a GT device as a range extender has been studied in previous works by the same authors (Capata and Sciubba 2009, 2010), where the most convenient matching of compressor and turbine was selected (under the assumption that the GT operates at fixed point) on the basis of a series of real mappings (courtesy of the Garrett/Honeywell R&D department), and the identified “best” operating point is reported in Fig. 7. Figure 8 shows how the two components match. The operating point for the compressor, corresponding to a corrected air flow of about 6 lb/min with a pressure ratio of two is located near the stall line at a velocity of about 175 000 rpm (to achieve maximum efficiency). The turbine operating point is located on the curve identified by a slightly lower pressure ratio (due to the pressure losses in the CC; Capata and Sciubba 2009), and by the properly increased mass flow rate (air + fuel). 3.1.1. The air preheater The air pre-heater and the CC have been conceived, designed, developed, patented (Capata and Sciubba 2011), and manufactured at the Mechanical and Aerospace Engineering Department of University La Sapienza-Roma, Rome, Italy. Their most relevant feature is the innovative single-body design: the air preheater spirals along the outer face of the cylindrical CC. The results of previous process simulations (Capata and Sciubba 2010) suggested the recourse to a preheating of the air upstream of Published by NRC Research Press



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Fig. 6. Load coefficient, Δcp, distribution on the wing: (a) α = 1°; and (b) α = 10°.

the combustor realized by inserting an additional duct that spirals around the CC outer wall, a solution that leads to a significant efficiency increase and reduces the danger of overheating downstream of the turbine, thanks also to the compressor relative value of absolute ΔT = (Tout – Tin) = 136 K. 3.1.2. Compressor The device is a modified Garrett GT14 radial compressor and its design and operation have been discussed in previous papers (Capata and Sciubba 2008, 2009). The number of rotor blades is 12: six full-span and six splitters, slightly backward-tilted at compressor exit (β2 = 62; α2 = 12), to maintain a rotor outlet velocity within limits dictated by the absence of a diffuser (some degree of pressure recovery takes place in the spiralling case that connects the compressor to the CC). The compression ratio is about 2/2.1 3.1.2.1. Tests measurements The experimental campaign was carried out according to the procedures described in detail in Capata and Sciubba (2009). In the following paragraphs, therefore, only the main results are described. Published by NRC Research Press


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Table 1. Cruise speed aerodynamic- and propulsive characteristics. CL 0.31

CD

Drag (N)

Power (W)

0.031

51.49

2476

Table 2. Dash speed aerodynamic and propulsive characteristics. CL 0.16

CD

Drag (N)

Power (W)

0.023

68.67

2636

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Fig. 7. Matching of the operating point on the (a) compressor and (b) turbine maps and (c) 3D rendering of proposed GT device.




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Fig. 8. (a) Computer-aided design rendering and (b) actual photo of the inner body of the combustor chamber with a detailed view of the preheater.

3.1.2.2. Compressor maps A dedicated test bench was built (Fig. 9) to reproduce the structural stress on the device under “cold” conditions (the turbine being driven by a jet of compressed air at room temperature) and to reconstruct the compressor map. A volumetric high-precision flow meter, a pressure gauge and a thermocouple are inserted on the turbine inlet channel to monitor the inlet conditions and to ensure steady operation. Similar devices are inserted on the compressor outlet channel, to measure the pressure ratio and to provide the data needed to compute the compressor efficiency. A thermocouple and a pressure gauge are inserted on the turbine exhaust. The results are summarized here in Figs. 10–13 and Table 3 3.1.3. CC As explained earlier, the CC is an original design patented (Capata and Sciubba 2011) by the Mechanical and Aerospace Engineering Department of the University Roma Sapienza. A first-order thermofluiddynamic simulation was performed to assess the attainment of a satisfactory thermal field and the completeness of combustion (Minotti and Sciubba 2010). A series of specifically dedicated Published by NRC Research Press


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Fig. 9. Functional sketch of the test bench.

tests, reported in Calabria et al. (2013) provided the temperature map of the CC for a more exact analysis of the thermal flows and for experimentally validating the computational fluid dynamics results. 3.1.3.1. CC thermal field A fundamental test for the CC is the reconstruction of the temperature map of its surface (Fig. 14). Measurements have been taken at the following points: . inlet compressor air in the preheating zone; . external wall of the inner cylinder; Published by NRC Research Press



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Fig. 10. The test bench of the UMGT.

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Fig. 11. The rotor of the UMGT tested (a 2 euro coin shown for comparison).

. external wall of the outer cylinder; . external wall of the exhaust gas volute (Fig. 15). The tests have thus far revealed that the stability range for this CC is rather small (which was expected on the basis of the results of the computational fluid dynamics simulations) and that the combustion efficiency is also very sensitive to the fuel/air ratio (Table 5). After the ignition protocol was satisfactorily validated, the next series of tests was devoted to the measurement of the temperature at points A, B, C, D, E, and F on the external wall of the chamber: . In section D (see Fig. 16), the values of temperature range from 1173 to 1288 K. These discrepancies appeared consistently throughout the tests, and therefore cannot be caused by the flow unsteadiness (though the chamber design produces a sort of pulsating combustion (Minotti and Sciubba 2010). . On the external chamber wall (surface C in Fig. 17), the measured temperatures varied between 473 K on the top cover (cold zone) and 873 K on the bottom one (hot zone). The cooler compressor air that flushes the preheating section exerts a moderating effect on the wall temperature. Published by NRC Research Press


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Fig. 12. The compressor rotor shape and velocity triangles.

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Fig. 13. Efficiency and outlet temperature versus compression ratio.

. A temperature measurement on the surface of the inner cylinder can be only performed in the immediate vicinity of the compressor air inlet section (point E, Fig. 16). The compressor air preheating obtained (a temperature increase of about 400 K) is reasonable in view of the following considerations: a. the temperature is relative to a wall on which “cold” (∼310 K) compressed air impinges; and b. the air “cup” temperature in the entry portion of the channel is lower than the average wall temperature of the inner cylinder. These measurements confirm that the CC attains a regeneration degree Rreg = (T2reg – T2)/(T4 – T2) equal to the one (0.8) assumed in the preliminary design phase. The results of the thermal test of the chamber are reported in Table 6. Published by NRC Research Press



Table 3. Compressor experimental data. pin, (atm) 1.002 1.002 1.014 1.015 1.025 1.027 1.030 1.035 1.039 1.048 1.089 1.091

Tout, (K)

βcomp

η

387 387 385 386 385 383 384 393 408 417 417 434

1.98 1.98 2.00 2.01 2.02 2.02 2.03 2.03 2.03 2.04 2.04 2.05

0.66 0.66 0.68 0.68 0.69 0.70 0.70 0.65 0.58 0.55 0.55 0.50

Table 4. Data acquisition time. Acquisition time

Start-up Transient operation Steady state

1–2 minutes 3–5 minutes 10–15 minutes

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Operational mode

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Fig. 14. The tested prototype of the combustion chamber.

3.1.4. The turbine The turbine used in this ultramicro turbo group is a Garrett/Honeywell® radial machine (Honeywell 2012), its body of cast Ni-based (Inconel) alloy with nine highly twisted blades. In line with usual construction practice, the gas has a purely radial inlet and an almost axial outlet velocity, to maximize the specific work. Published by NRC Research Press


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Table 5. Ignition time. CC conditions

Ignition time

Cold Warm

30 s Instantaneous

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Fig. 15. Sketch of the flows in the CC.

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3.1.5. The dual effect electric engine Packaging constraints have suggested in this phase to adopt a configuration with a single shaft for compressor and turbine: the alternative solution with a double shaft, that splits the power required by the compressor from that delivered by the turbine, and leads to a more complicated arrangement of electrical devices. The double-shaft configuration is worthy of future investigation though, because it would reach a higher efficiency because of the possibility of independently selecting the optimal rotational range for each machine. One of the advantages of the single shaft configuration is that it has an intrinsically well-balanced volume–weight distribution and that it automatically ensures matching between the rotors. The switching of the electrical drive motor to an electrical generator when the turbine generates enough power to drive the compressor is controlled by an ad hoc electronic device that measures the amperometric absorption of the engine. Unfortunately, the electrical efficiency of this dual effect (motor– generator) device is lower than that of its single effect equivalent. The currently installed electrical engine–generator has been designed and manufactured by an Italian company (Nuova C.I.S., Busto Arsizio, Italy) and is capable of reaching an angular velocity of 20 933 rad/s (200 000 rpm). A lower speed of 170 000, however, is sufficient to allow the turbine to drive the compressor during the start-up phase and to reach a self-sustaining regime. The electrical device is passively cooled by external cooling fins, and its total installed weight is 5 N. 3.1.6. The management unit The management unit performs several tasks related to the control and monitoring of the correct behaviour of every component at “design” operation and during the start-up and shut-down phases. This is the actual brain of the turbo group, ensuring the shutting off of the fuel injection in case of overheating or any other malfunction. The control is performed through the inlet gas valve on the basis of instantaneous data obtained from the thermocouples monitoring the temperature and Published by NRC Research Press



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Fig. 16. Thermal field of UMGT UDR1 combustion chamber.

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Fig. 17. Assembled configuration dimensions and rendering (dimensions in mm).

pressure of the main flows. The management unit has been adapted, by proper reprogramming, from a commercially available prewired electronic board. 3.1.7. Size and weight The adopted configuration has the axis of the GT group placed parallel to the axis of the CC. This arrangement results in a very compact envelope without affecting the access to any component both in the installation and in the maintenance phase. The dimensions of the assembled device are shown in Fig. 17 (the supporting fixture is the one used in the laboratory, and a different one will be developed for flight tests). Published by NRC Research Press


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Table 6. Combustion chamber thermal field. Position A B C D E F

Temperature (K)

Top cover Bottom cover External surface Exhaust gases Inlet air duct* External outlet duct*

473 873 523–673 1173–1288 523 873

*After a 5 min warm-up period.

Table 7. Overall dimensions of the UMGT device, proposed configuration. L (mm) 184

b (mm)

h (mm)

334

190

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Fig. 18. The UMGT prototype mounted on the test bench.

The weights of the components of the proposed configuration have been estimated, and are reported in Table 7: these values (still approximated and probably about 25% higher than the final engineered solution) serve as an indication of maximum total weight of the GT group, necessary information for the calculation of the payload of the UAV (Fig. 18; Table 8).

4. Conclusion In this paper a conceptual design of a high-endurance UAV, the preliminary design of hybrid propulsion. The drone has a flying wing configuration with a twin tail and an aft-propeller located in the rear centre. The aerodynamic analysis was aimed at the investigation of the lift and drag characteristics of the drone and to a preliminary calculation of its thrust and power requirement to fulfil the mission specifics. The choice of a flying wing leads to a reduction of the wetted surface and of the overall drag, Published by NRC Research Press



Table 8. Approximate component weights, total ∼46 N. Component

Material

CC Compressor (including volute) Turbine (including volute) Oil pump Gearbox Supporting structures Electrical motor and generator

Stainless steel Steel Steel, hastelloy rotor Steel Steel Aluminium Steel, copper, iron

Weight (N) 25 4 4 1 2 5 5

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improving the performance of the UAV. The volume available to fit the payload box and the propulsion plant was verified and a layout of the components is given. The UMGT design satisfies all design specifications (payload and size). The feasibility of the project has been validated and future activities will consist of the construction of a mock-up of the drone and of a flying prototype for performance tests. System and experimental tests on its prototypal version have been presented and discussed. Recalling that under steady flight conditions thrust is equal to the resistance and that in the cruise operating mode we need T = 51.5 N and at maximum speed T = 68.7 N at the propeller shaft (the electric motor drives the propeller) the required power must be increased to account for the efficiency of the propeller (approximately 0.7), which results in a nominal power request of 2476 and 2636 W, respectively. The system was conceived so that the energy flux from the GT device is used to recharge the battery package and not to directly provide power to the propeller. However for safety reasons the turbine power is equal to that of the electric motor in case of a failure or malfunction of battery package. But because this is only an emergency operational mode, an On–Off operating mode has been evaluated here. As a final consideration, it must be noticed that, with current technology, it is possible to recharge the Li–ion battery package with an electrical current up to six times higher than the maximum allowable operating current (6 C). Hence, the adoption of a recharging mode of 2 or 3 C adopted in this study can be seen as a way of extending the life of the battery package. The preliminary bench tests indicate a specific fuel consumption of about 840 g/kWh (corresponding to an overall efficiency of about 6%). These performance data may appear dismal if compared to standard-size engines (for example, the Capstone 45 kW GT declares a specific fuel onsumption of 250 g/kWh), but it must be kept in mind that a reduced engine size leads to a substantial decrease of the efficiency, for GT but also (even if less strikingly so) for ICE (a Honda 1500 cc automobile engine declares an optimal-point specific fuel consumption of 250 g/kWh, while the smaller 5 kW for aircraft model applications has a specific fuel consumption of about 700 g/kWh, the famous Maynard Hill 0.75 kW transatlantic model aircraft being a noteworthy exception with a specific fuel consumption of 406 g/kWh). The performance of the two propulsion systems (GT-hybrid versus ICE) ought to be evaluated on a specific mission basis: assuming a 12 h mission, and that both the propulsive power and the specific fuel consumption remain constant throughout, an Exdrone-like (Ronney 2005; Barnard 2006) ICE-propelled plane would have to carry approximately 420 N of fuel and the GT-powered drone described in this work (with the GT operating 1/3 of the time), 60 N. The respective weights would 1 be 150 N for the ICE engine, about 50 N for the GT and 300 N for the Li–ion battery pack: although the advantage of the GT-hybrid is apparent, its lower-than-specification payload (20 N) indicates that more work is needed in three areas: (i) the weight of the GT-set; (ii) its specific fuel consumption; and (iii) the battery weight. In view of the other operational advantages of a GT-hybrid concept, all of these area are worthy of further investigation.

References Barnard Microsystems Ltd. 2006. UAV Microsystems. Available from http://www.barnardmicrosystems.com/about/papers_ &_presentations.html [accessed April 2013]. Bertin, J., and Smith, M. 1998. Aerodynamics for Engineers, 3rd Ed., Prentice Hall. Calabria, A., Capata, R., Di Veroli, M., and Pepe, G. 2013. Testing of the Ultra-Micro Gas Turbine Devices (1–10 kW) for Portable Power Generation at University of Roma 1: First Tests Results. Eng. 5: 481–489. doi: 10.4236/eng.2013.55058 Capata, R., and Sciubba, E. 2006. Preliminary considerations on the thermodynamics feasibility and possible designs of Ultra-, Micro- and Nano-gas turbines. Int. J. Thermodynamics. 9(2): 81–91. doi: 10.5541/ijot-308 1. The weight of the UMGT does not scale linearly with power. This figure, based on the values reported in Table 2, is to be considered an educated guess.



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Capata, R., and Sciubba, E. 2008. Progress in the development of a prototype “Ultra-Micro” Gas Turbine Set for Portable Power Generation. In Proc. of ECOS 2008 Conference, Volume II, p. 573–580, Gliwice, Poland. Capata, R., and Sciubba, E. 2009. The Ultra-Micro Gas Turbine Generator project at UDR1: experimental assessment of the compressor map and of the regenerative combustion chamber efficiency. In Proc of ECOS 2009, p. 2069–2078 (ISSN 2175-5426) Foz Do Iguaçù, Brasil. Capata, R., Marino, L., and Sciubba, E. 2010. Preliminary design of a hybrid propulsion system for high- endurance UAV. In Proc of International Mechanical Engineering Conference & Exhibition IMECE2010-38530, p. 107–112; doi:10.1115/IMECE2010-38530, ISBN: 978-0-7918-4425-0, 2010 Vancouver, Canada. Capata, R., and Sciubba, E. 2011. [Patent]. Improved Turbogas: micro regenerative combustion chamber for ultra micro gas turbine devices. Italian Patent ID RM2011 A 000277(03/06/2011). Chao, H., Luo, Y., Di, L., and Chen, Y.Q. 2010. Roll-channel fractional order controller design for a small fixed wing unmanned aerial vehicle. Contr. Eng. Pract. 18: 761–772. Gur, O., and Rosen, A. 2009a. A Design of a Quiet Propeller for and Electric Mini Unmanned Aerial Vehicles. J. Propul. Power. 25: 717, doi: 10.2514/1.43733 Gur, O., and Rosen, A. 2009b. Optimizing Electric Propulsion Systems for Unmanned Aerial Vehicles. J. Aircraft. 46: 1340, doi: 10.2514/1.C031020 Himoinsa power generation motors. 2014. Available from http://www.himoinsa.com/Product/Product.aspx [accessed September 2013]. Hoerner, S.F. 1965. Fluid-Dynamic Drag. Hoerner Fluid Dynamics, Bricktown, New Jersey. Honeywell. Garrett turbobags. 2012. Available from http:www.turbobygarrett.com [accessed December 2011] Kim, J., Van, N.N., Lee, J., and Kim. S. 2012. Integrated Design and Analysis for Electric-Powered Unmanned Aerial Vehicle Optimization. In Proceedings of 12th AIAA Aviation Technology, Integration, and Operations (ATIO) Conference and 14th AIAA/ISSM 17–19 September 2012, AIAA 2012-5609, Indianapolis, Indiana. Logan, J.M., Chu, J., Motter, A.M., Carter, L.D., Ol, M., and Zeune, C. 2007. Small UAV Research and Evolution in Long Endurance Electric Powered Vehicles. In Proceedings of AIAA Infotech @ Aerospace 2007 Conference and Exhibit, AIAA Paper 2007–2730, 7–10 May, 2007. McCormick, B.W. 1995. Aerodynamics, Aeronautics and Flight Mechanics. Wiley & Sons. McGhee, R.J., and Beasly, W. 1973. Low-Speed Aerodynamic Characteristics of 17 percent thick airfoil section designed for general aviation application. NASA Technical Note TN-D-7428. National Aeronautics and Space Administration Langley Research Center, Hampton, VA, December 1973. Available from http://www.ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19740003708_1974003708.pdf Minotti, A., and Sciubba, E. 2010. LES of a meso combustion chamber with detailed Chemistry Model: comparison between the flamelet and EDC models. Energies. 2010(3): 1943–1959. doi: 10.3390/en3121943 Moran, J. 1984. Computational Fluid Dynamics. Wiley & Sons. Paterson, J.H., MacWilkinson, D.G., and Blackerby, W.T. 1973. A Survey of Drag Prediction Techniques Applicable to Subsonic and Transonic Aircraft Design. Aerodynamic Drag, AGARD-CP-124, Neuilly-sur-Seine, France, October 1973. pp. 1-1–1-38. Ronney, P.D. 2005. Hydrocarbon-fueled internal combustion engines: “the worst form of vehicle propulsion … except for all the other forms”. Available from ronney.usc.edu/whyicengines/WhyICEngines.pdf [accessed January 2014]. Roskam, J. 1985. Airplane Design. Vol. VI. Roskam Aviation and Engineering Corp. Schoemann, J., and Hornung, M. 2012. Modeling of Hybrid-Electric Propulsion Systems for Small Unmanned Aerial Vehicle. In Proceedings of 12th AIAA Aviation Technology, Integration, and Operations (ATIO) Conference and 14th AIAA/ISSM 17–19 September 2012, AIAA 2012-5610, Indianapolis, Indiana. Teknel Power Generation Motors. 2014. Available from http://www.teknel.net/ [accessed September 2013]. U.S. Department of Defense. 2001. Unmanned Aerial Vehicles Roadmap 2000–2025. Available from http://www.globalsecurity.org/ intell/library/reports/2001/uavr0401.htm#_Toc509906768 [accessed August 2013]. Vleugels, P., Waumans, T., Peirs, J., Al-Bender, F., and Reynaerts, D. 2006. High-speed bearings for micro Gas turbines: stability analysis of foil bearings. J. Micromech. Micro-engineering. 16(9): 282–289. doi: 10.1088/0960-1317.

List of symbols A a b bt CD CL cr ct e FF h l S T Tin Tout V W α2 β

Aspect ratio Aspect ratio Span Span at tail Drag coefficient Lift coefficient Chord at wing root Chord at wing tip Oswald factor Form Factor Altitude Length Wing surface (m2) Temperature (K), Resistance (N) Compressor inlet temperature (K) Compressor outlet temperature (K) Relative wind velocity (m/s) UAV weight (N) Physical blade angle (absolute; °) Compression ratio



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Physical blade angle (relative; °) Leading edge sweep angle (°) Taper ratio Air density (kg/m3)

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β2 Λ λ ρ



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