adapting a small ground turbine to very light airborne applications: a “lean” approach

8/12/2019 Adapting a Small Ground Turbine to Very Light Airborne Applications: a Lean Approach

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Presented at the American Helicopter Society 64th Annual Forum, Montreal, Quebec, Canada, April 29 th May 1st, 2008. Copyright 2008 bythe American Helicopter Society International, Inc. All rights reserved.

Adapting a Small Ground Turbineto Very Light Airborne Applications:

a Lean Approach

Mauro Piazzoli and Teddi RossiAviotecnica - Mantova - Italy

[email protected]

Giovanni Gasparotto [email protected]

Giuseppe Gasparini [email protected]

Abstract

There is a need of practical and real alternatives to the more common gasoline internal combustion engines burningthe increasingly difficult to find Avgas fuel, in particular for the Very Light Airborne applications, and this isexpected to increase even further in the next future.This paper describes the authors efforts in research and experimental activity to verify the feasibility of theadaptation of an existing turbine for ground based, power generation use to Very Light Airborne applications,

primarily Very Light Rotorcrafts (VLR), UAV and Very Light Aircrafts (VLA).The selected starting point was an existing SOLAR T-62T-32 ground based, power generation gas turbine.The SOLAR T-62T-32 was originally used as prime mover for ground based, generator sets. It is of the single shaftdesign, single stage centrifugal compressor, reverse annular combustor, single stage radial in-flow turbine, so it is ofvery simple and rugged design even if particular care has to be paid to its assembly and tuning to achieve acceptable

performances, efficiency and controllability in the new airborne environment.The team initial work involved CFD simulation of the compressor-diffuser assembly to identify critical performance

parameters and subsequent rig (Dyno) testing to correlate analytical predictions with actual data to fine tune theanalytical models and to establish a performance baseline to be used later on to verify and assess the possibleimprovements.Then, a complete review of all the external and internal mechanical parts was performed.The objective was to achieve an output shaft power of about 130 HP at S.L. ISA at 6000 RPM (corresponding to thestriking speed of about 61,091 RPM at the turbine mainshaft!) but at lower weight and with increased reliability.The weight reduction was achieved by detail work on turbine inlet case, gear cases, gears and accessoriessubstantiated by stress analysis, gear and bearing calculations, rotor dynamic studies and the definition of specificoperational limitations to avoid exceedance of unacceptable stresses.In addition, a dedicated Electronic Engine Control (EEC) has been designed, manufactured and tested to allow therepresentative operation of the modified engine.Extensive bench testing (in excess of 300 running hours on the lead modified turbine) confirmed the consistentachievement of the above mentioned objectives and lead to the experimental installation of a modified turbine on aAviotecnica RAVEN prototype (AUW about 450 Kg, two seats VLR) where it has been flight tested.


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1 Introduction

The need of practical and real alternatives to the morecommon gasoline internal combustion engines burningthe increasingly difficult to find Avgas fuel, is quiteevident in particular for the very light airborneapplications, and it is expected to increase even further

in the next future.As an example of available engines for very lightaircraft, it can be mentioned the internal combustionengine ROTAX 914 which is in the range of 115 HP,

but is prescribed not for rotary wing use in the officialInstallation Manual and has a weight of about 7274 Kg(Ref. 1).The big aerospace companies working on UAV seem to

be also looking at the diesel engine or at the gas turbine, but it appears that there is still a lack of readily availablesolutions for the very light range of applications(100 150 hp).

The authors have combined their passions for aviationand for gas turbines to perform a research andexperimental activity to verify the feasibility of theadaptation of an existing turbine for ground based,

power generation use to very light airborne applications, primarily Very Light Rotorcrafts (VLR), UAV and VeryLight Aircrafts (VLA).

Being the budget for this research extremely limited and based mainly on the personal resources and time of theteam members, it was governed by very stringent lean

principles: always consider the real essence of any partor activity, avoid any redundancy or waste, consider and

prefer always the simplest, cheapest and more easy tomanufacture, install and maintain solution, etc.

The selected starting point was an existing SOLAR T-62T-32 ground based, power generation gas turbine.The choice was simply dictated by the followingconsiderations:

Simple design, low part count, ruggedconstruction

Easy availability of used turbines in fairlyacceptable conditions

Affordable prices for used units (servicesurplus)

The SOLAR T-62T-32 was originally used as primemover for ground based, generator sets. It is of thesingle shaft design, single stage centrifugal compressor,reverse flow annular combustor, single stage radial in-flow turbine, so it is of very simple and rugged design

even if particular care has to be paid to its assembly andtuning to achieve acceptable performances, efficiency andcontrollability in the new airborne environment.It has also the advantage of burning conventional, widespread turbine engine fuels (i.e. Jet A1).

From the very early stage of the project it became clearthat the three main directions of investigation had to befollowed:

Performance and Efficiency review andimprovement of the turbomachinery

Structural Integrity of the mechanical parts

Weight Reduction of the overall engine

2 Description

The solar T-62T-32 Gas Turbine (Fig. 2.1), also calledTitan, was originally designed as a ground based powergeneration set.

Fig. 2.1 The Solar T-62T-32 Gas Turbine

Considering the turbomachinery part of the generator set,it is a simple gas turbine of the single shaft type with

single stage centrifugal compressor, reverse flow annularcombustor with six burners and single stage radial inflowturbine.

The nominal speed of the turbine shaft is 61091 rpmwhich is reduced to about 6000 rpm by a star type(rotating sun and ring gear, fixed planet carrier) epicyclicreduction gear stage.


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The gearbox housing the main reduction gearing alsodrive the drive train for the accessories (pumps,generator, de-aerator).

The design is very simple and rugged and minimize the part count and the complexity so reducing cost andincreasing reliability making this engine one goodchoice for low cost applications.

One drawback of this engine concept is the fact that itneeds a clutch to start rotorcraft rotors providing aappreciable resistance during starting. This is becausethe turbine delivers very small power during the starting

phase, just sufficient to drive the compressor.

For this reason, it is necessary to equip the engine with aclutch, normally of the centrifugal type, such that theload (i.e. the rotorcraft rotors) is engaged only when theengine has reached a speed where it can deliver powerin excess of the power needed to drive the compressor(self sustaining speed): in this particular caseAviotecnica designed and developed a centrifugal clutchengaging at about 50% rpm, but this will not bedescribed in this paper since it is considered outside thescope of this paper.

Table IT-62T-32 Main Characteristics (Ref. 2)

Rated Speed 61091 rpmRated Power 90 hpGearbox Output Speed 6000 rpmWeight (dry) 76 KgMaximum exhaust gas temperature 1180 FSFC at Rated load, 60F 0.64 Kg/hp-hr

The original characteristics of the legacy engine arereported in Table I. It was judged by the authors that theoriginal configuration was not acceptable for airborneuse from the point of view of efficiency, weight,

performance and structural integrity.

3 Turbomachinery Analysis

3. 1 Introduction

In order to increase the efficiency of the turbine enginein the cheapest and easiest possible ways, a study ofoverall of compressor performance was needed (Ref. 3).Compressor is deemed as the most critical part in a gasturbine due to the complex internal flow path. Howeverit was considered relatively simple to increaseefficiency through an optimization of the compressordesign.

3. 2 Reverse Engineering

The first activity undertaken was the reverse engineeringof the legacy compressor wheel and diffuser vanes.This was achieved by inspecting an existing wheel on aDEA 5 axis Coordinate Measuring Machine (CMM).The cloud of points obtained was then processed bymeans of the SurfCAM CAD-CAM software to obtain 3Dsurfaces (Fig. 3.1).

Fig. 3.1 Reverse Engineered 3D Model ofCompressor-Diffuser Assembly

3. 3 CFD Analysis Preprocessing

The above mentioned 3D geometry was used to define the

mesh of the Computational fluid Dynamics (CFD) model.

Fig. 3.2 Diffuser Mesh


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The solid elements mesh was then obtained with thehelp of Fluents GAMBIT mesher (Ref. 4), Fig. 3.2 and3.3.

Fig. 3.3 Impeller Mesh

Some iterations were needed to obtain an acceptablemesh with tetrahedron elements.

3.4 CFD Analysis Solution

The actual CFD analysis was performed using theFulents solver (Version 5.4.8) available at theMechanical Engineering Department of the Universityof Padua, Italy.The non-stationary option of the solver could have

permitted to more realistically simulate in details thestator-rotor interactions, but this option was not adopteddue to the unfavorable ratio between the impellernumber of blades and the number of blades of the radialcascade of the diffuser which would have involved themodeling of the entire compressor with unacceptablecomputational costs (more than 2 millions cells).For this reason, the model was adapted and simplified to

be compatible with the actual computational resourcesavailable: a stationary model with separate simulation ofthe impeller and of the diffuser.It was therefore sufficient to mesh a segment of theimpeller (one eighth) and one segment (one third) of thediffuser obtaining, respectively 130,000 cells and200,000 cells.The turbulence model adopted was the standard k- model.The analysis was defined in order to obtain thecharacteristics curves for five rotational speed at 105,

100, 90, 80, 70 % of the nominal speed (i.e. 61,091 rpm).

3. 5 CFD Analysis Results

The complete operating maps of the compressor impellerwere obtained.

Fig. 3.4a CFD Results for the Impeller

Fig. 3.4b CFD Results for the Impeller

Fig. 3.4c CFD Results for the Impeller


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For the impeller, it can be noted that at the nominaloperating point the inlet velocity field is totally subsonicand at the outlet it is almost transonic.On the other hand, the diffuser is characterized by asupersonic velocity field in the throat of the diffuserat the operating point, determining at that point thechocking of the entire compressor (Fig. 3.5a and3.5b).

Fig. 3.5a CFD Results for the Diffuser

Fig. 3.5b CFD Results for the Diffuser

The analyzed velocity field appears quite widesometime inside the unstable operation characterized bythe surge phenomenon.

A further extension of the analysis to extrapolate thecurves would not have been possible due to theimpossibility of the solver to converge within a domaincharacterized by both an area of instability and bysupersonic waves.

The results of the analytical work are brieflysummarized in Fig. 3.6a and 3.6b but they will bediscussed in more detail in the next chapter.

2

2,5

3

3,5

4

4,5

5

5,5

0,6 0,65 0,7 0,75 0,8 0,85 0,9 0,95 1 1,05

Corrected Mass Flow Rate [kg/s]

C o m p r e s s

i o n

R a

t i o

100%

90%

80%

70%

Ne

Fig. 3.6a Calculated Compressor Characteristics(Compression Ratio)

0,7

0,75

0,8

0,85

0,9

0,95

1

0,65 0,7 0,75 0,8 0,85 0,9 0,95 1 1,05

Corrected Mass Flow Rate [kg/s]

I s e n

t r o p

i c E f f i c i e n c y

10090%80

70%

Neq.

Fig. 3.6b Calculated Compressor Characteristics(Isentropic Efficiency)

4 Testing

A test campaign was undertaken with the followingobjectives:

To validate the numerical simulation in

accordance with the requirements of ASMEPTC 10-1997 (Ref. 5).

To assess the performance behavior of the engineunder different conditions

The complete operating maps of the compressor impellerwere obtained.


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Fig. 4.1 The T-62T-32 on the Aviotecnica Test Rig

4. 1 Instrumentation

According to the above mentioned standard, 4 points ofmeasurement have to be taken for each parameter on 2measuring stations to obtain the operating maps.In addition, other probes have been added at thecompressor impeller outlet.

In summary, we had the following measurement points:

Compressor Inlet:

4 ports for total pressure connected to a waterdifferential pressure transducer

4 ports for static pressure connected to a waterdifferential pressure transducer

4 probes for total temperature (type Kthermocouples)

1 pitot tube to measure the mass flow rateconnected to a differential water pressuretransducer

Compressor Outlet (Fig. 4.2):

4 ports for total pressure connected to a waterdifferential pressure transducer

Fig. 4.2 Compressor Outlet Pressure Ports

Diffuser Outlet (Fig. 4.3):

4 ports for static pressure connected to a mercurydifferential pressure transducer

Fig. 4.3 Diffuser Outlet Pressure Ports

The engine instrumented as described above was theninstalled in the Aviotecnica facility engine test cell (aDyno, Borghi & Saveri FE450-SV) for full speed and

power testing (Fig. 4.1 and 4.4).


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Fig. 4.4 The Instrumented T-62T-32

4.2. Experimental Results and Comparison withAnalytical Predictions

The characteristic curves of the overall compressor,obtained both experimentally and numerically, areshown in Fig. 4.5.

Fig. 4.5 Comparison Between Experimental andComputed Compressor Characteristics (Ref.6).

The pressure ratio p03/p01 and isentropic efficiency are plotted as functions of the corrected mass flow rate atfour values of the rotational speed: n, 0.9n, 0.8n and0.7n.The experimental test showed that the compressor hasquite a narrow operating range at all rotational speeds,and that the normal working point is very close to thechoke line, the corrected mass flow being 1.033 kg/sand the pressure ratio 3.6. In this condition, theisentropic efficiency is about 0.70.This operating point presumably gives adequate marginsagainst the occurrence of compressor surge withoutheavy drawbacks on the efficiency. At reducedrotational speeds, the pressure ratio curves flatten out

and suggest how careful the operation in these conditionsshould be in order to avoid the occurrence of flowinstabilities.In some cases, significant differences were registered inthe code computing accuracy concerning pressure ratioand efficiency.At nominal rotational speed, predictions of the pressureratio are in excellent agreement over the whole operatingrange, while those regarding the efficiency are lessaccurate, the maximum discrepancy being in the order of4%.At reduced rotational speeds, computed values of theefficiency are quantitatively better while those of pressureratio are worse (i.e. the code underestimates the pressurerise).In fact, as the rotational speed reduces, the computedcharacteristic curves are shifted toward lower values ofthe mass flow rate, i.e. the code found some difficulties incapturing the choke condition. While examining theseresults, however, the uncertainty on experimental data aswell as the limitations of the steady state approach and ofthe turbulence model, especially at part load, must be

properly taken into account: the use of the steady-stateapproach, in particular, is known to give misleadingresults when the operating conditions are very far fromthe nominal one and strong recirculation at impeller-diffuser interface are usually observed. In this case,however, the use of a mixing plane still gives acceptableresults because of the narrow operating range of thecompressor.Also, the authors believe that some of the discrepanciescould be explained with the absence of the tip clearance inthe numerical simulations, which would limit the

maximum mass flow rate to some extents and wouldcontribute to a reduction in the total pressure ratio andefficiency.The characteristic curves of the diffuser are reported inFig. 4.6

Fig. 4.6 Comparison of Test vs Calculated Data(Diffuser Characteristics, Ref.6)


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They show both the experimental and computed valuesof the pressure recovery coefficient Cp and theaerodynamic loss coefficient as functions of thecorrected mass flow rate.At the nominal speed, the measured pressure recoverycoefficient decreases from 0.5 to 0.45 as the mass flow

rate increases from stall to choke (i.e. as the angle of theabsolute velocity with respect the tangential directionincreases from 19.4 to 19.9). In such conditions, theaerodynamic loss coefficient increases from 0.45 to0.51.The same tendency is qualitatively observed at reducedspeeds, even though the pressure recovery coefficient ishigher at part load, because the dynamic pressure atimpeller exit decreases much more than the diffuser

pressure recovery. On the other hand, the aerodynamicloss coefficient reduces slightly at lower mass flow ratesdue to the fast decrement of the total pressure losswithin the diffuser.Because of the lack of measurement probes in theannular bend, it was not possible to isolate the effect ofthe radial and axial blade rows on overall diffuser

performance and, therefore, to establish whether or notthe axial de-swirl cascade has a negative effect oncompressor efficiency and stability. The agreement

between the numerical and experimental results is goodat design nominal speed, whereas areas of relatively

poor code accuracy were found at reduced rotationalspeed. Again, this can be justified if the simplificationsof the numerical model at impellerdiffuser interfaceare taken into account.In any case, the overall results suggest that the CFDmodel is sufficiently accurate to give realistic

indications on diffuser performance within the overallcompressor operating range.

4.3 Possible Optimizations

The natural consequence of the analytical andexperimental work on the existing, legacy configuration,was the authors decision to assess the potential foroptimization of this engine concept.

A numerical multi-objective optimization of the radialand de-swirl cascades of the centrifugal compressor was

accomplished through an iterative procedure based onthe combination between a Multi-ObjectiveEvolutionary Algorithm (MOEA) and a CFD model ofthe diffuser (see Ref. 6 for more details).The aim is to develop a set of diffuser designs achievingmaximum pressure rise (maximum Cp) and minimumtotal pressure loss (maximum 1- ) at the designcondition. These designs have also to fit into the radial

and axial sizes of the original one, perhaps adjusting theradius between the radial and the de-swirl sections.

The most important design parameters and thecorresponding objective function values for the lastgeneration of solution are presented in Fig. 4.7. Note that

performance indexes of the original design are too low(Cp = 0.45 and 1- = 0.5) to appear in the figure.

Fig. 4.7 Results of Diffuser Optimization(Ref. 6)

Thanks to the optimization of compressor design (up toPressure Ratio of 6), turbine new materials (max.temperature increase up to 914C) and the results of ourtests a projected, optimized configuration was defined asshown in Tab. II.

Table IIComparison of Existing and

Optimized (Projected) Turbines

T-62T-32Existing

T-61BProjected

80 Hp 135 Hp 80 Hp 135 Hp

Pressure Ratio 3.6 6.0TIT [C] 660 814 744 910ThermalEfficiency [%]

11 14 16.5 20

FuelConsumption[Lit/hr]

60 83 42 61

5 Mechanical Design Review

The mechanical review of the existing hardware (Fig. 5.1)was aimed at:


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Verifying its suitability for the airborneapplication

Improving the mechanical safety

Reducing the weight

Fig. 5.1 Mechanical Parts to be Modified

Due to simplicity of the original turbine design, itclearly became apparent that the most important andvital mechanical part of turbine was the rotor assembly,therefore most of the activities concentrated on thisassembly.

5. 1 The Rotor Assembly

The core of the engine is its rotor assembly (Fig. 5.2).The rotor assembly is basically composed of:

The mainshaft assembly (including the nut and pins)

The front ball bearing

The rear roller bearing

The compressor impeller

The turbine impeller

The bearing cartridge

Fig. 5.2 Rotor Assembly(Schematic)

This assembly is a very critical part of the turbine enginedue to the high speed of rotation which requires extremedimensional and geometrical accuracy, careful assembly,

tight dynamic balancing.

5. 2 Rotor Dynamics Study

The very high rotational speed of the rotor assemblycombined with the significant mass of rotors and theslenderness of the mainshaft dictated as the first priorityfor our mechanical investigation to study the dynamical

behavior of this assembly starting with evaluation of itscritical speeds.The significant mass and polar inertia of the rotors withrespect to the stiffness of the mainshaft clearly suggestedthat without considering the gyroscopic effects of therotating masses, the critical speeds would have beenincorrectly predicted.

The basic theory for the mathematical analysis of thiskind of problem is based on the following equation:

[ ] [ ] [ ] { } { }012 =+

K Gh

M

where:

[M] : concentrated (lumped) mass matrix[G] : gyroscopic matrix[K] : stiffness matrix{} : eigenvector (mode of vibration)

==GYRA

h1

gyroscopic factor

1+=GYRA forward synchronous precession(forward whirl)


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0=GYRA no gyroscopic effect

1=GYRA backward synchronous precession(backward whirl)

[ ]

=

pn

p

J

J

G

0

00

..

..0

00

1

Gyroscopic matrix

The analysis, in practice, has to be performed by digitalcomputing with general purpose finite element method(FEM) software (Ref. 7) or with rotor-dynamics

software (Ref. 8).

Fig. 5.3 The Gyroscopic Effect

The results can be presented in a way that isqualitatively represented on Figure 5.3 in order toidentify the most potentially dangerous critical speedsof the rotor assembly. This graphical representation,although different from those of the above mentionedreferences, is preferred by the authors for its moreimmediate and compact description of the phenomenain a single quadrant and its consistency with othervibration problems (e.g. rotating disks resonance).

For our study the most affordable approach was to usethe simplified calculation method reduced to 2 DOF (i.e.the radial and angular displacement of the impellersCG),

As confirmed by Ref. 9, the forward whirl (forward precession) speeds are the most dangerous because theyare excited by the unbalances always present in a

rotating structure (for this reason, also, the obsessive carein balancing rotating parts which will be described lateron is justified).In the case of forward whirl, the gyroscopic moment hasthe tendency to increase the stiffness of the shaft, henceincrease its natural frequency.Very simply speaking, we can think about the gyroscopiceffect as a tendency of a high inertia rotor (the impellersin this case) to maintain its rotational axis and to resist toany attempt to change the orientation of the axis ofrotation.

From the critical speed analysis above, see Table III, it isconcluded that the mainshaft is actually supercritical (i.e.it normally operates above its first critical speed) but hasgood margins when operated at nominal speed and also atthe Ground Idle (GI) speeds.

Table IIICalculated Critical Speeds

These results bring some considerations:

In a possible future redesign or optimization of aturbine of this type it should be preferable tointroduce some form of non-rotating damping inorder to reduce the vibration during the crossingof the critical speed and to avoid any possibilityof sub-synchronous instability at nominal speed.This could be achieved, for example, byintroducing squeeze film damping on the roller

bearing.

Experimental, precise assessment of the criticalspeed placement should be performed on the rigto confirm the analytical prediction (this has notyet been performed due to unavailability at

Aviotecnica of the required vibration analysisequipment.

5. 3 Main-shaft

In addition to the very obvious need of higher reliability,a mechanical system to function in airborne conditions

Mode Description Speed

rpm

Percent

%

1 Shaft bending 12900 21

2 Shaft bending coupled withroller bearing radial

249000 407

CriticalSpeed


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has to be verified also for the loads induced by the flightmaneuvers.

Fig. 5.4 Main Shaft Assembly

In particular for high speed rotating machinery the loadsinduced by gyroscopic effect cannot be neglected.As it was seen earlier, the gyroscopic effect influencessignificantly the dynamic behavior of the rotorassembly, in this case it has also a significant effect onthe strength of the mainshaft (Fig. 5.4).In fact, due to the small size of the shaft compared tothe high polar inertia of the impellers (compressor andturbine) connected to it, the gyroscopic moment inducedstresses have to be considered.More specifically, a bending moment has to beevaluated as follows:

C=J*( m * elic )whereC : gyroscopic momentJ : polar moment of inertia of the rotor

assemblym : angular speed of the rotor assemblyelic : angular speed of aircraft (yaw

maneuver)By stress analysis of the mainshaft under the combinedloading condition it was concluded that the most criticalsection was the shaft section with the lubrication holes(Fig. 5.5 ).In that section we have the combined effect of the

gyroscopic bending moment, the transmitted torque andthe unbalance forces, in addition we have unfavorablestress concentration factors, both in bending and intorsion (and in this section we have unfortunately both

bending and torsion) due to the radial holes (Ref. 10).

Several ideas and studies were evaluated in order todefine a solution to improve the mainshaft strength at itscritical section, but they all involved a significantredesign and manufacturing of this complex part whichwas not compatible with the budget and schedule of this

project.Therefore it was decided to introduce a specificlimitation on the use of this engine in flight in order toavoid the risk of overstressing of this shaft section.

Fig. 5.5 The Critical Section of the Main Shaft

In conclusion it was imposed:

elic 2 rad/sec

The respect of this limitation will assure that the stress atthe critical section of the mainshaft will remain belowadequate limits.

5. 4 Gears

The main reduction gearing were evaluated and wereconsidered suitable for the intended use in airborneapplications.The only changes were:

Reduction of weight by machining of theexisting parts (Fig. 5.6)

Dynamic balancing

Replacement of the ball bearings of the planetswith roller bearings to increase load capacity andreduce power losses

Fig. 5.6 Gears Lightening

CRITICALSECTION


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5. 5 Bearings

The gyroscopic moments affecting the shaft strengthwill inevitably stress also the bearings, therefore theirsizing should be verified accordingly.

For the above reason (i.e. to cope with the additionalgyroscopic loads) and to increase their reliability, the

bearings were completely redesigned in collaborationwith an aerospace bearing manufacturer (SNFA, nowSKF Aeroengine, France).

For the calculation of the bearing lives a simplified dutycycle (power spectrum) was established as reportedon Table IV.

Table IV

N. CONDITION POWER

%1 TAKE-OFF AND LANDING 1002 HOVERING IGE + OGE 803 FORWARD FLIGHT (CRUISE) 754 FORWARD FLIGHT (V-MAX) 805 MANEOUVER 1106 DESCENT AND LANDING 60

MAINSHAFT DUTY CYCLEFOR BEARING CALCULATION S

5

100

61091

10

451015

6109161091610916109161091

15

DURATION SPEED

% rpm

For the ball bearing (Fig. 5.7) having the function to provide front radial support as well as thrust reaction, itwas decided to adopt a conventional 3 contact points(split inner ring) made of aerospace quality M50 toolsteel, with outer land riding cage made of silver platedsteel of high accuracy (ABEC 7).Outer land riding was selected because of thelubricating condition of this bearing: it is lubricated bythe oil mist coming laterally from the roller bearing.

Fig. 5.7 Ball Bearing

The rear roller bearing (5.7) presented a more substantialchallenge due to the peculiar aspects of this application.In fact, even if the roller bearing has to react the radialload generated by the unbalance of the rotors and by themoments induced by the gyroscopic effects, most of itslife is spent with very low radial loads applied.This, combined with the high rotational speed (dN >10 6),results in a significant risk of skidding (inner race slippingrelative to the rollers).

The two most practical measures to mitigate this risk wereidentified as:

Minimizing the operating radial clearance of theroller bearing

Designing the cage inner land riding such thatthe shaft drive the cage

The decision to design the cage inner land riding prompted the review of the lubrication of this bearing and, particularly, of its cage.

Fig. 5.8 Bearings Lubrication Scheme

The legacy turbine engine lubrication system had to bemaintained for practical reasons, and it is schematicallydepicted on Figure 5.8: oil is injected into the sun gear

bore where it flow axially and finally escape throughradial holes on the back of the roller bearing by means of

centrifugal force.

Oil is then contained on the right side by a labyrinth sealand drawn to the left by the depression induced by aradial rotating impeller (the slinger) on the left of the

ball bearing.

Nut/Slinger

Labyrinth

Ball Brg Roller Brg

Sun Gear


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Fig. 5.9 Roller Bearing

This lubrication system feeding the roller bearing fromthe inside was considered suitable for the inner landriding cage and suggested to improve the bearing design

by adopting the philosophy of cage design described inRef. 11 whereas the cage is shaped with wingletscatching the oil and directing the catched oil to thecritical contact inside the bearing.This concept was finally adopted and developed incollaboration with SNFA.

5. 6 Gearcases

Due to the extremely limited budget and manpoweravailable, it was not possible to implement the mostobvious design improvement of the gearbox casings:that is the change from cast aluminum alloy tomagnesium even if this change alone would have saved6 to 7 kg.However, in order to reduce weight without losingstrength of the casings, they were submitted to adetailed machine and hand finishing to remove excessmaterial, to improve finish and to smooth fillets (Fig.5.10)

Fig. 5.10 Modified Gearbox Casings

Nevertheless, it is very clear that for future development,a more drastic change of these components has to bemade with the adoption of a high quality magnesiumalloy, such as for example the more recent Elektron 21(Ref. 12) which combines high static and fatigue strength,higher corrosion resistance, good high temperatureresistance and acceptable castability.

5. 7 Dynamic Balancing

Due to the extremely high speed of the rotor and thedynamic considerations explained under therotordynamics chapter, it can be concluded that the

precise and accurate dynamic balancing of all the rotating parts is of extreme importance for the proper and safeoperation of this engine (Fig. 5.11).

Increased reliability was pursued by systematic andobsessive care of assembly and dynamic balancing of allrotating parts.

Fig. 5.11 Rotor Assembly Being Balanced

Based on authors experience it can be concluded that themain factors affecting the optimum dynamic balance are:

Extremely careful design of balancing arbors andfixture

Suitable drive for spinning the rotors withoutgenerating parasitic vibrations

Rigid assembly balancing sequence involvingindividual parts balancing to Grade G2.5 or

better according to ISO1940 (Ref. 13), assemblycentering checks and assembly balancing againto G2.5 or better


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5. 8 Weight Reduction

As in every real aerospace project, the weight is themost important factor and several Weight ReductionCampaigns have to be undertaken during thedevelopment.In the case of this project, several improvements withrespect to the original hardware were introduced asdescribed in this paper to reduce the final mass of theengine.

The most significant ones were: Redesign of the generator Redesign of the EEC Modification of the gears Modification of the casings

At the end the total weight of the T-61A engine wasreduced to 63 Kg (about 12 Kg less than the original T-62T-32!).

6 Engine Controls

The existing engine controls of the legacy ground basedgas turbine were judged as not suitable for any airborneapplication for both weight and performanceconsiderations.Therefore it was necessary to work hardly on bothmodifying the existing hardware but also in designingand developing new components.

6.1 Engine Fuel Control

The hydro-mechanical components of the engine fuelcontrol were substantially maintained as the originalones, except for minor improvements to the fuel pump,extremely accurate adjustment and tuning of the systemand for the addition of a fuel accumulator (Fig. 6.1 and6.2).The basic purpose of the fuel accumulator is to allow asmooth transition from starting mode to full powermode and vice-versa. More specifically, this permit toachieve the followings:

Avoid engine shut down when passing fromfull power (Flight, 100% rpm) to start(Ground Idle, 30% rpm) by keeping

pressurized the manifold circuit whichotherwise would reduce sharply its pressurewhen the start circuit opens

Avoid engine shut down in case of engineoverspeed (exceeding 115% rpm). In this casethe fuel control would automatically select the

start circuit to reduce speed. As before, thiswould reduce too rapidly the pressure in themain circuit with the risk of engine shutdown

Damp the power burst when passing from thestarting mode to the full power mode

Fig. 6.1 Fuel Accumulator Cross Section(Schematic)

The device is installed on the engine gearbox and is provided with a spring loaded piston accumulating pressurized fuel which is then released when the pressurein the main circuit decreases.

Fig. 6.2 The Installed Fuel Accumulator

6.2 Electronic Engine Control

The Electronic Engine Control (EEC) was completelyredesigned in order to reduce the size and weight of theequipment and to tune its characteristics to those requiredfor the airborne application (Fig. 6.3).

Vent

Fuel inlet fromfuel controlassembl

To themanifold

circuit


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Fig. 6.3 The New (left) Vs. the Old (Right) EEC

The EEC was designed using analog electronics (Fig.6.4) due to the limited time and budget available but,nevertheless, it achieved more than acceptable

performance when properly adjusted.It receives the input signal from the engine speedmagnetic pick-up and regulate the fuel flow to maintainthe speed of the engine constant independently of thetorque demanded.The redesign also contributed significantly to the weightreductions with more than one Kg of saving.

Fig. 6.4 Inside the New EEC

6.3 Starter-Generator

The existing starter fitted on the turbine was clearlyunsuitable for the intended use. In fact it was very heavyand lacking of the function of generator (Fig. 6.5).It was necessary to completely replace it with a newone.The only practical and affordable solution was to adoptthe rotor of an existing motorbike starter-generator anddesign around it a suitable stator mating with the

existing drive pad on the engine gearbox (Fig. 6.5 and6.6).

Fig. 6.5 The Old Starter

In addition, the redesign offered the opportunity toimprove the cooling of the unit by equipping it with an aircooling system exploiting a small fraction of the enginecompressor airflow to induce an air cooling flow through

the starter-generator fitted with screened air intake andcarbon fiber collector.

Fig. 6.6a Cross Section of the New Starter Generator

Fig. 6.6b The Assembled New Starter-Generator

The weight saving for this component case wassignificant with about 5 Kg of reduction and an addedfunctionality.

MotorbikeRotor


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7 Ground and Flight Testing

The finally configured modified turbine engine,identified as T-61A (Fig. 7.1) , has been quiteextensively tested both on the rig (Dyno) and on a

flying prototype of the Aviotecnica RAVEN two seatssport helicopter kit.

Fig. 7.1 The T-61A Turboshaft Ready for Test

The main characteristics of the finally configuredturbine are as listed on Table V.

Table VT-61A Main Characteristics

Rated Speed 61091 rpmMax. continuous power 90 hpMaximum EGT @ MCP 915 FTake-Off power 130 hpMaximum EGT @ TOP 1180 FGearbox Output Speed 6000 rpmSelf Sustaining Speed 20%Ground Idle Speed 30 %Clutch Engagement Speed 50 %

Nominal Operating Speed 100%Maximum Overspeed 115%Weight (dry) 63 KgFuel consumption at MCP, 60F 62 L/hSFC at MCP, 60F 0.496 Kg/hp-hr

The performance characteristics and weight of theengine are considered suitable for a rotorcraft of GrossWeight between 450 and 650 Kg.

Ground testing was performed on the same conventionalautomotive test rig (Dyno) of the Aviotecnica test

facility (Fig. 7.2) which was used previously for performance mapping and characterization and was wellcompatible with the shaft horsepower of the turbine.

Fig. 7.2 The T-61A on the Test Rig of AviotecnicaTest Cell Facility

In addition to the performance assessment, the groundtesting has been also a confirmation of the very good levelof dynamic balance and assembly accuracy achieved.In fact, as a routine practice, the vibration level on theturbine casing is always measured by means of aChadwick-Helmut (now Honeywell) Vibrex analyzerand provided very low vibration levels at all the testedconditions.

Extensive flight testing was performed on a protype of theAviotecnica Raven kit for 2 seats sport helicopter (Fig.7.3).

Fig. 7.3 The Aviotecnica RAVEN VLR prototype with themodified turbineT-61A installed during flight testing

The T-61A always behaved satisfactorily through theflight test (now about 300 hours have been accumulated


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on a single engine) confirming good performance andreliability and excellent response during maneuvering.8 Perspectives

The analytical and experimental work and theencouraging results obtained with the modified engineT-61A stimulated the investigation about the potential

perspectives of this engine concept.On Table VI a first comparison of T-61A against

current gasoline internal combustion engines is presented.

Table VIComparison of Current Engines for VLR/VLA/UAV

EngineModel

T-61A Rotax914

LycomingO-360

Type Turbine 4 cylindersIC

4 cylinders IC

Fuel Jet A1 Avgas orUnleaded

Avgas

Rated Power 130 hp 115 hp 180 hpOutputSpeed

6000 rpm 2300 rpm 2700 rpm

Weight* 63 kg 75 kg 118 kgFuelConsumption@ 75% P.

62 L/h 20.4 L/h 40 L/h

SFC @ 75%P.

0.458Kg/h/hp

0.170Kg/h/hp

0.213Kg/h/hp

* to the turbine engine an additional reduction from 6000 rpm toabout 2500 has to be added, but, on the other hand, to thereciprocating engine the cooling system (radiator, fan, water or fanand shrouds) has to be added, it is assumed that they are equivalent

and therefore they cancel out in this comparison

This comparison shows that, despite the muchlower weight if the turbine engine and its use of betterfuels it has higher SFC than the gasoline enginesmaking its use more expensive.

However, in its power range, the performanceadvantage, low vibration, fast maneuver response stillmake its choice, the most attractive option.

But considering the most advanced perspective ofadvanced turbine (considering the potential

performance, efficiency and weight improvements ofchapter 4) and diesel internal combustion engines (Ref.14) the comparison summarized on Table VII can bemade.

In this case the T-61B is the projected turbineobtained introducing all the performance, weight andefficiency improvements identified.

This table shows that the two solutions with same

fuel (Jet A1) and same power still show a SFC advantageof the internal combustion engine but much less than

before and with a weight penalty for the diesel engine ofabout 77 Kg.

Table VIIComparison of Advanced Engines for VLR/VLA/UAV

Engine Model T-61B DieselType Turbine 4 cylinders

Diesel

Fuel Jet A1 Jet A1Rated Power 135 hp 135 hpOutput Speed 6000 rpm 2300 rpmWeight 57 kg * 135 kgFuel Consumption@ 75% P.

55 L/h 21.5 L/h

SFC @ 75% P. 0.391Kg/h/hp

0.179Kg/h/hp

TBO 2000 hrs 2400 hrs* 6 kg of saving are estimated from the adoption of magnesium

in place of aluminium for the geracases

Matching the two advanced propulsion solutions ofTable VII to an hypothetical UAV rotorcraft of about 630Kg gross weight yields the performance resultssummarized on Table VIII.

Table VIIIComparison of UAV Rotorcraft Performance

Engine Model T-61B DieselRated Power 135 hp 135 hpEngine Weight 57 Kg 134 KgGross Weight 630 Kg 630 KgFuel Weight 98 Kg 21 KgFuel Volume 136 L 29 LEmpty Weight * 295 Kg 295 KgPayload 180 Kg 180 KgRange 2.5 Hrs 1.2 Hrs

* excluding engine

It is clearly shown that for this specific application an

optimized turbine is still largely superior in terms of range(more than two times the diesel equipped UAV!) inaddition to the faster response and controllability and lowvibration advantages.


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9 Conclusions

In conclusion, the research study described in this paper has demonstrated that a small single shaft, singlestage gas turbine with proper adaptation to the airborneuse like the T-61A or, better, its optimized projectedevolution, the T-61B, are suitable solutions for

powering VLR, VLA and UAV airborne vehicles in therange of 450 to 650 Kg AUW, providing:

Affordable costs

High reliability

Mechanical simplicity

Low vibration level (no reciprocating masses)

Fast response/controllability to pilot inputs

High power to weight ratios

Readily available fuels

It is also believed that an improved or newlydeveloped, simple, single shaft, single stage, radial flowgas turbine, like the optimized projected evolution T-61B, could benefit significantly from state-of-the-arttechnology in turbomachinery design, materials, etc.

providing the increased performance and efficiencywhich are needed for a large scale application on lighthelicopters, airplanes and UAVs.

To achieve these development objectives it is deemedthat, based on the work already performed atAviotecnica, further steps have to be done.In particular, from the point of view of performance andefficiency, the following areas should be developed:

new turbomachinery (in particular thecompressor-diffuser) with improvedaerodynamics to increase efficiency

improved internal air sealing (replacement oflabyrinth with brush seals)

optimized air intake and exhaust system

new digital engine control

new gearbox casing in magnesium to reduceweight by about 6 kg

In addition, to increase the structural integrity and themechanical reliability, the following areas should bedeveloped:

new mainshaft to increase strength and fatiguelife

introduction of squeeze film damper for theroller bearing to reduce the vibrations, in

particular during the critical speed crossing

10 References

1. Installation Manual for ROTAX Engine Type914 Series Bombardier ROTAX, 2006

2. T-62T-32 Technical Manual Overhaul GasTurbine Assembly Air Force 1989

3. Giovanni Gasparotto, Microturbine per UsoAeronautico Studio delle Prestazioni delCompressore ( Microturbines for AersopaceUse Compressor Performances Study ) -Engineering Graduation Thesis, The Universityof Padua, Italy, 2002 (Unpublished)

4. Fluent 5 Users Guide - Fluent Inc.

5. ASME PTC 10-1997: Performance Test codeon Compressor and Exhauster

6. E. Benini, A. Toffolo, A. Lazzaretto,Experimental and Numerical Analyses toEnhance the Performance of a MicroturbineDiffuser Experimental Thermal and FluidScience, 2005

7. H. Bedrossian, N. Veikos, Rotor Disk SystemGyroscopic Effect in MSC/Nastran DynamicSolutions - Avco Lycoming Division

8. G. Genta, Vibration of Structures andMachines Springer-Verlag, 1995

9. T. Krysinski, F. Malburet, MechanicalVibrations Active an Passive Control ISTE2007


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10. W.D. Pilkey, Petersons stress concentrationFactors 2 nd Ed. J. Whiley & Sons, 1997

11. Anonymous, Aeroengine Roller BearingDesign - RollsRoyce

12. P. Lyon, ELEKTRON 21 A NewMagnesium Alloy Developed for Aerospaceand Specialty Applications MagnesiumElektron, 2003

13. H. Schneider, Balancing Technology Schenck 1991

14. M. Rozenkranc, J. Ernst, Diesel Engines forAerospace Use IAI, 2003

adapting a small ground turbine to very light airborne applications: a “lean” approach

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