a mathematical model and experimental investigation on increasing starting altitude of turbine...

8/3/2019 A Mathematical Model and Experimental Investigation on Increasing Starting Altitude of Turbine Engine

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A M A 'f IEMAl ' ICS MOIIEI, AND EXPERIMEN'T 'AI,JNVES7'lGATION ON INCREASING STARTINGAL,'TlPUllfi OF TURBINE ENGINE

Ni Xing-Qiang and I , i n W en.[et Propulsion Depai-trneiit

Beijing, P R ChinaI3eijing University of Aeronautic s and Astro nautics

A theoretical and experimental investigation has beencarried ou t to study tlie effect of oxy gen addition on th ealtitude ignition perform ance of a tui-bine engine Th eresults show that tlie injection of oxygen caii improvealtitude light-up performa nce significantly. .]'he startingaltitude of turbine engine ha s been increased approximatelyfi-om 4000 meters to 8000 meters. T he light-up flight Machnuniber has been varied roughly from 0.36 to 0 8

I n the theoretical phase of this investigation, the paperhas calculated the flow structure of a shoit-annularreverse-flow combustion chamber of a turbine engine. Thecombustor created vortices to stabilize combustion byusing three jet groups. The SIM PLE (Semi-ImplicitMethod for Pressures-Linked Equation ) and constantviscosity turbulent model were em ployed to solve theNavier-Stokes equations. The paper has predicted the flowstruct ure The flow field of calculation sho ws that thereverse flow vortex at the back of the air inlet tube is morestable Here can form a combustion zone for he1 addition.

In the experimental phase of this investigation. the basictheory of increasing starting altitude ha s been discussed indetails. An air- borne oxygen supply ignition device( AOSID ) has been developed to expand the applied rangeof the engi ne. High altitude simulating cell tests and flighttests with m other air-craft or turbine engine equipped withAOSID have also been performed successhll\i.

Subseciuently. tlie results of ignition tests show thatAOSlD is very beneficial to ignition, especially to highaltitude ignition.

INTIIODUC'IION

The turbine engine in this paper w as originally employedTh e turbine engine is a small one-sliatt jet engine. The

in the rem ote pilot vehicles.fuel is injected from the shaA to the annular combustionchambe r Th e ignition system consists o f star-ling file1nozzles and high energy sparking plugs ( Fig. I ). The

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flanie tiley made goes into the head of flanie tube, and thenlights the main fuel injected by the fuel distributor

Figure I Th e coinbustion chamber of the originalturbine cngineThe iwwiniuni ignition altitude ofthe turbine engine

without oxygen addition is only 4000 meters to 5000meters. and the maxiinurn ignition Mach nuniber is only0 36. l h e I-equirenient oft he parent aircraft engine is 7000meters and Mach 0 55 or 9000 meters and Mach 0 8respectively These are the obstacles for the turbine enginein use So , it is very important for the engine to im provealtitude ignition performance

'The factors of which the altitude ignition for e iy n e isunder influence are very complicated -1hei-eare nianymethods to improve i t One of thein is ail-borne oxygensupply ignition device ( AOSID ) Based on our theoreticaland expcriinental investigation for inany years in BeijirigUniversity of Aeronautics and Astronautics, it is a capablemethod i n resolving the problenis of altitude ignitionperformance

M A T H E M A I I C s MODEL

The turbine engine ignition is a very complex physicaland chemical period. It includes three phases which areboth relative and dependent

The first phase, the electric spark plugs discharge tobuild up a flanie kernel which is big eno ugh and itstempera ture is high enou gh. These make it possible forflame to propagate

The sec ond phase, ii'thei-e is a comb ustion which isbigger and occupies partially in the combustion chambe r,there must he a f low stiuctiire ( i.e., reverse-flow region )

CH34827-9510000-28.1 $1.00"'1995 EEE


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where

Semi-Implicit Method for Pressure-Linked Equation( SI M PL E ) has been used in this paper- to solute thetraiisrornied go verning equations until a steady-state solutionis reached

l. inally, the pi-ograni n this paper rail on ItlM-4341 inEleijing liniversity of Aeronautics and Astronautics I tincludes 10 subroutines It iterated 150 steps every time 'Therunning-time of CPU was 30 minutes

The mathematics model of the turbine engine ha s beenobtained on the basis of the calculation 1-esults Fig. 2)

'The results show there are three jet groups in combustionchamber which were supplied by the compressor.

'The first group is injected by the louvers Its peripheralvelocity i s 15 i d s , 10 m i s , 5 rnis respectively

The second group is injected by the air inlet tube Its radialvelocity i s 10 ni/sThe third gr oup is injected by the inner combuster shell. Itsvelocity of every hole is 25 i d s .

I t can be seen clearly in Fig. 2that there are three vorticesin the flow field

The first reverse vortex is the biggest main flow vortex I tis in the center of the chamber head I t is the determinant for

Figure 2 The grid system and flow field of annularcombustion chamber

creating a stable cornbustion. It is necessary to light it duringthe altitude ignition. The third group combined w ith the firstgroup at the chamber wall, and then turned backward. Themixture group were intercepted by the second group at the airinlet tube to form the first reverse vortex,

According to many tests i n which we used the tempera tureindicating paint and the 'Titanium powd er p aste to visualize

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to stabilize the flame. and then. it can become a morepowri- ignition source ,in the combustion chamber in order to make thecombustion expand to all the main fuel The combustio nchamber works normally with it Now. th e ignition periodi s over.I t is impossible for us to solve all probleins about the allignition period of combustion cham ber only withtheoretical calculation meth od. It involves the calculationof three-dimension visco us flow, two-phase flow.thermocheinical problems and etc. We think it is possiblefor the turbine engine developed an air-bo rne oxygensiipply ignition device ( AOSID ) to start the altitudeignition systeni reliably due to investigations and analyseson the basic physical conceptio ns of ignition andcombustionobtained as fcdlows.

7'hc third phase, the ignited flames ignite the other fuel

Flow m odel of annular combustion chaniber has been

1 Governing EquationsIhe calculation in this paper used two-dimension, steady,

incompressible, axisymnietric Navier-Stok es equation(Reference i , 2 )

( 1 ) Continuity

( 2 ) hlomentum

v v e v a " 0

whereVr , V ,, and Ve are velocity components in the radial, axial

and peripheral directionsan d

v = vc -t- MvF Kinetic molecular viscosity coefficientvt Kinetic turbulent viscosity coefficientFor constant viscosity turbulent models, it is thought that

the visc0sii.y coefficients are equal everywhere in full flowfield. When the turbulent fluctu ation isvery violent in thecombu stion chamber, the turbulence level is 20 percent to 40percent. It is thought that the actual viscosity coefficient is100 to 200 times as big as the viscosity coefficient.

2 Solution of the Governing EquationsBecause the chamber wall is irregular, that i s to say, it has

curved boundary, TT M coordinate conversion was used inthis paper to aenera te automatically body-fitted grid system.'This is practicable in dealing with the complicated curvedboundary of the combustor and viscosity fluid reversevortices. saving the conipute r memoi-ies and times. In order tosolve the Navier-S tokes equations with this grid system, thegoverning equations must be transformed froni circularcylindrical coordinate to curvilinear coordinate After applyingthe transformation, the governing equations can expressed as

( I ) Continuity

where

(2) Momentum

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the flow structure, this reverse-flow region is at the directionof 45 angle in the head of the flame tubetlie chamber head The first group flowed along the wall,encountered the third group, and was ended The blendedgroup turned back to produce tlie inferior vortex. ~1 e vortexis smaller than the main vortex. We th ink that i t existed in th ecalculation. but in fact. it did not existbuilt up by the induction of the jet flow injected by the air inlettube.

The line of boundary between the first group and the thirdgroup can be seen clearly in Fig 2 .

The flow structure, i e , he positions and dimensions of thevortices, are different because of the different parameters.such as tlie distribution of inlet air of the jet gro up, ttievelocity. the injected angle of tlie jet groups. and therotational speed of the engine, etc Tlie calculation has beenmade under various conditions. Subsequently, the results ariddiagrams show that the reverse flow at the back of the air iiileltube is more stable Here can form a comblistion zone for fueladdition which can be ignited directly

The second vortex is an inferior vortex I t is at the below of

The third vortex is at the back of the air inlet tube It is

S C H E M E OF IMPROVEh5ENT

T he lo w starting altitude and velocity of the original turbineengine are produced by tlie characteristics of the ignitionsystem designation. It s starting nozzles and electric sparkingplugs are performed at the flow tunnel between the radialdiffuser and flame tub e case Tlie ignition method is tunneltorch. The torch goes into tlie inner conibustor shell in theflame tube to ignite the main fuel and to become a stable flamein the reverse flow region. When the rotational windmillingspeed o f he turbine engine is 700 rpmiinin to 800 rpmimin atthe altitude of 4OOO meters to 5000 meters and Mach 0 36The flow velocity near the starting fuel nozzles is about lessthan 12 mis. So, the flame source is stable. When tlie altitudeincreases to 7000-9000 meters and the Mach number is morethan 0.55 to 0.8, the rotational windmilling speed is 2000r p d m i n to 2500 rpmiinin The flow velocity Rear the nozzlesis about more than 40 niis to 60 ni/s 1-lie eniperalure i ncombustion chamber is -27 centigrade or so N o w , t hestarting flame source is unstable Clearly, the disadvantag es ofthis ignition method i s that the low temperature & pressure.less oxygen and high windmilling speed make the startingflame unstable th e tunnel and difficult in ignition when theflight altitude and velocity increased. In addition, the hightempe rature flame make easily the radial diffuser out of shapeI t is true to make the engine burned because of it before

Certified by the theoretical and experimental investigations,oxygen addition in the altitude can increase the amount o foxygen i n the mixture air, velocity ofreleas ing heat and flam e

temperature. decrease the minimum ignition energy. improvethe ignition performance Hence , it is a practical inethod inincreasing starting altitude.

I he criteria about increasing altitude ignition of the engiricwere carried out based on the results of he above calculatioiland man> tests

I 7-here must be a stable flame source2 The ignition must be in a stable vortex of atomized fuel

in the cornbustion chamber.3 . When the am ount of oxygen is less i n the altitude, the

oxygen additioii is very necessary. The position of the device( i \OSlD) can not be in the original ignition position

According to the ct-iteria above, the ignition device u a sselected in this paper as followed.

I About tlie selection of the stable flanie sourceThe ignitmn statement of the oiiginal turbine engine\+.e selected the indirect ignition method which was uszd in

the other engines. i e , pre-chamber ignition (Fig 3 ) Th eignition w as ver-y stable This kind igniter can start to workreliably a t the designation statement i n which ttie altitude is12000 meters, tlie pressure in igniter is 0 22 kg/clnz , th e lloi4velocity is 4000 niis to 5000 mis , he temperature is -3 0cent gi-ade

Figure 3 Indirect starting igniter

2 'The ignition po sition i n the main vortex of atomizedfile1

The igriition position i n the original engin e is between theradial difiser and the flame tube ( I t i s called "turn path" forshort ). 'The main fiiels are ignited there. During the altitudeignition, it can not be lighted-up due to !he unstable startingflame source Somebody used the method of the oxygenaddition in the turn path. As a result. the engine w as burnedout for tile reasons of tlie structure and etc.

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Considering the above, w ith the indirect ignition niethod,the holes were made in the flame tube to create a stablestarting flame to ignite the chamber. 7-his method can resolvestarting problems of the main reverse flow region in altitudeignition. We redesigned the ignition system in order to ignitethe main reverse flow region with the starting flame morereliably.

Proved by the tests, our iinprovernent is successfiil

increased several hundred ce ntigrade. The ignition ability ofthe starting flame and altitude ip it io n capability have beenanielioratcd

4 The s(:heme ofthe improving starting altitude ignition ._.A O S I D

In accordance with the above analyses, we used ,A.OS[D(Fig 4)

Figure 4 The position of the new igniter

3 . About the utility of oxygen additionThere is not AOSlD in original engine. To develop th e

AOSID can increase the amount of oxygen in the mixture air.the releasing heat velocity and modify the ignitionperformance because there is low pressure and low density inhigh altitude. On the basis of ou r designation, the amount o foxygen addition in the igniter is 0 .9 g/s to 1 . 1 ds, he oxygencoefficient an increased from 0.21 to 0 .3 2-4 .35 in thecondition of the standard atmosphere. The ignition capabilitycan be improved after the oxy gen addition as follows.

( I ) Reducing the ignition energy(2) Increasing the pressure in the igniter. This can incre ase

the pressure from 0.22 kg/cm z to 0 32 k&m* .

(3 ) . ModitLing the excess air coefficient in tlie igniterThe original excess air coefficient is 0.12Afker oxygen addition, the excess air coeflicient is 0 286(4). ?'he oxygen exponen t after oxygen addition increa sed

from 0.21 to 0 . 32 .The temperature of the starting flame

I R E X ,

The tests were performed a t the ground test rig, the highaltitude simulating cell and the flight cell respectively Followsare the results

1 . The ground testThree engines ( J69-T-4 1A,B,C ) were tested for 270 times

and we got a number of test data These are the reliable datato assure the starting altitude ignition fluidly. Fig 5shows themodel curves i n tlie ground test

I ------ Blade tip temperature2 ------ Fuel flow3 Rotational speed4 Flame signal

Figure 5 The model curves irn the gi-ound test


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2 The high altitude simulating cell ( HASC' ) testl ' wo engines ( 569-1'4 A, B ) were tested for 20 times at

th e t 1ASC The altitude is 4000 me ters. 600 meters. 9000meters respectively The blach nuniber is 0 4 to 0 8 -Jheignition for engine equipped with AOSID have beenperfoi me d successfully

Fig 6 s h o~ . she H A S C test of the engine with AOSIDl'ig 7 shows the inodel test cui-ves of th e 36Y-T-41B.

Figure 6 I IASC' test of the engine with AOSI D

Figure 7 The model c urves of J-69-T-4 IB i n I lASC test( Altitude 9250 meters, Mach 0 8 7 )

3 The flight cell test'The J69-T -4 113 engine was tested for 7 flights and 20 times

ignition te sts i n the high altitude. The altitude is 9000 meters,7000 meters, 4000 nieters, 1900 meters respectively. TheMach number is 0 4 to 0 8 The test results show the turbineengine can be ignited reliably during starting period in aboveconditions This niakes the turbine engine tit for the throwingdemand of unmanned aircrafts and missiles, and expends theapplied rang e of it

Fig 8 s h o w the flight envelope line of the mother aircraftFig 9 show s the flight test of the J69-T-41B with AOSIDFig IO shows the flight test curve s.

3 i2: ; 2 0 0 4 0 5 0 6 0 7 0 6 '''

Figure 8 Flight envelope of the mother aircraft

Figure 9 Flight test of the J69-T-41 B with AOSID

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CONCLUSION

I The altitude ignition ofturbine engine is a technicaldifficulty The investigations of increasing starting altitudemake it successful for the utility of the turbine engine

2. The factors influencing the altitude ignition are verycomplicated. I t is impossible for us to resolute the problemsonly with theoretical method Combined with experimentalinvestigation, the altitude ignition perforitlance can beincreased.

Ignition Altitude 9000 niFlight Velocity 500 km/hMach Number 0.712Windmilling Speed 233 1 rpndminAtniosphere 'Temperature - I 5.3 centigrade

Optical Discharge ( volt )REFERENCE

1 Ning-hang, " Combustor Aerodynamics ", eijing,P R China, 1982

Exhuast Temperature ( centigrade )

2 Xu-Hung, " Coinputer Simulation RC ExperimentalVisualization of the Vortices in an Short-AnnularTurning-Flow Combustor ",Beijing University ofAeronautics & Astronautics, I982

3 Zhao Qi-Shou "Calculation of WindmillingCharacteristics of Turbojet Engine " ASME 80-GT-50

4 Private Communication

Pressure of Oxygen Addition ( KYa )120xn10

0

-1

Figure 10 Th e flight test curves

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