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Numerical Heat Transfer, Part A:Applications: An International Journal ofComputation and MethodologyPublication details, including instructions for authors andsubscription information:http://www.tandfonline.com/loi/unht20

The Effects of Heat Shielding in JetEngine Exhaust Systems on AircraftSurvivabilityKyung Joo Yi a , Seung Wook Baek a , Man Young Kim b , Sung NamLee c & Won Cheol Kim da Department of Aerospace Engineering , Korea Advanced Instituteof Science and Technology , Daejeon , Republic of Koreab Department of Aerospace Engineering , Chonbuk NationalUniversity , Jeonju , Chonbuk , Republic of Koreac Department of Nuclear Hydrogen Development andDemonstration , Korea Atomic Energy Research Institute , Daejeon ,Republic of Koread Agency for Defense Development , The 7th R&D Institute - 2 ,Daejeon , Republic of KoreaPublished online: 03 Apr 2014.

To cite this article: Kyung Joo Yi , Seung Wook Baek , Man Young Kim , Sung Nam Lee & Won CheolKim (2014) The Effects of Heat Shielding in Jet Engine Exhaust Systems on Aircraft Survivability,Numerical Heat Transfer, Part A: Applications: An International Journal of Computation andMethodology, 66:1, 89-106, DOI: 10.1080/10407782.2013.869441

To link to this article: http://dx.doi.org/10.1080/10407782.2013.869441

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Numerical Heat Transfer, Part A, 66: 89–106, 2014Copyright © Taylor & Francis Group, LLCISSN: 1040-7782 print/1521-0634 onlineDOI: 10.1080/10407782.2013.869441

THE EFFECTS OF HEAT SHIELDING IN JET ENGINEEXHAUST SYSTEMS ON AIRCRAFT SURVIVABILITY

Kyung Joo Yi1, Seung Wook Baek1, Man Young Kim2,Sung Nam Lee3, and Won Cheol Kim4

1Department of Aerospace Engineering, Korea Advanced Institute of Scienceand Technology, Daejeon, Republic of Korea2Department of Aerospace Engineering, Chonbuk National University, Jeonju,Chonbuk, Republic of Korea3Department of Nuclear Hydrogen Development and Demonstration, KoreaAtomic Energy Research Institute, Daejeon, Republic of Korea4The 7th R&D Institute - 2, Agency for Defense Development, Daejeon,Republic of Korea

The infrared signatures from hot engine parts pose major threats to military aircraftsurvivability. Reducing the skin temperature at the rear of the fuselage is key to reducingsusceptibility to heat-seeking armaments. A heat shield placed between the nozzle walland the outer casing of the engine can decrease the skin temperature at the rear of thefuselage. Therefore, numerical modeling of the fluid flow fields coupled with the radiativeand conductive processes within the heat shield, nozzle, and casing were performed todetermine the temperature distribution at the rear of the fuselage. The effect of thematerial properties and the dimensions of the heat shield were studied in order todetermine their effects on the susceptibility of an aircraft.

1. INTRODUCTION

Technologies associated with aircraft survivability and susceptibility tomilitary armaments are critical factors in modern warfare. Technological advantagedetermines supremacy in the modern battlefield and, consequently, research intosurvivability is extremely important. Defense contractors are engaged in equippingfighter planes with stealth capabilities in the early stages of the design process inorder to reduce susceptibility. For example, the advanced stealth technologiesemployed in Boeing’s Phantom Ray and Northrop Grumman’s X-47B unmannedcombat vehicles demonstrate the growing importance of reduced susceptibility ofunmanned combat vehicles. The primary threats to aircraft survivability result fromdetection using radar and thermal infrared (IR) sensing. A number of detection

Received 28 August 2013; accepted 7 November 2013.Address correspondence to Seung Wook Baek, Department of Aerospace Engineering, Korea

Advanced Institute of Science and Technology, 291 Daehak-ro, Yuseong-gu, Daejeon 305-701, Republicof Korea; E-mail: swbaek@kaist.ac.kr

89

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90 K. J. YI ET AL.

NOMENCLATURE

a speed of sound T temperatureaP coefficient of the discretized RTE u� v velocities

at nodal point P Ur reference velocity�A��V surface area and volume of the x� y Cartesian coordinate system

control volume Yi mass fractionDci directional weight � axisymmetric coefficiente total energy � pre-conditioning matrixE� F inviscid fluxes vectors � polar angleEv� Fv viscous fluxes vectors �a absorption coefficienth total enthalpy � densityHs axi-symmetric vector � viscous stress tensorI radiative intensity azimuthal anglek turbulence kinetic energy turbulence dissipationm mass flow rate � solid angleM Mach numberp pressure Superscriptsq heat flux m�n polar and azimuthal directionsQv pre-conditioned primitive

variables vector Subscripts�r position vector of b black body

radiative intensity c conductionR gas constant L�R left and right states of the�s direction vector of control volume face

radiative intensity r radiationt time 1/2 value at the control volume face

systems have been developed in recent years, and this technology has becomeincreasingly sensitive. Therefore, reducing the thermal and radar signatures ofaircraft is an important area of technological development. Stealth technologiesare effective in reducing the radar signatures of aircraft; however, losses fromheat-seeking missiles are a significant issue affecting aircraft survivability, and soanalyzing and reducing the IR signature of military aircraft is of critical importance.

Militarily advanced countries lead the survivability and susceptibility research;however this research is not disseminated and the outflow of these technologiesis strongly regulated because they are directly connected to national securityMahulikar et al. [1] illustrated the importance of infrared signature suppressionsystems and countermeasure techniques, and demonstrated the importance of theatmosphere in IR signature analysis. Research in the field of IR countermeasures(IRCMs)—devices designed to protect aircraft from IR seeking weapons—can becategorized into two classes: active and passive [1, 2]. Active countermeasures includejammers and flares dispensed by an aircraft to confuse the IR guidance system ofheat-seeking missiles by emitting dummy IR signals. Passive measures, termed IRsuppression focus on the nozzle geometry of jet engines and the flow mixing ofexhaust gases to minimize the IR signature from an aircraft. In turbofan engines,the IR plume signature can be reduced by increasing the aspect ratio of the nozzle[3]. Additionally, lobed forced mixers can suppress the IR signal from turbofan

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HEAT SHIELDING IN JET ENGINE EXHAUST SYSTEMS 91

engines [4]. The IR signature is strongly related to the skin temperature of theexhaust system; the temperature distribution at the rear of the fuselage of fighteraircraft can be used to estimate the IR signature levels [5]. Radiative heat transfercalculations using statistical narrow-band models have been used to determine thewall temperature distribution of aircraft exhaust systems [6], and the effect of thecharacteristics of the rear part of the fuselage on improving the survivability hasbeen examined [7]. IR signature suppression analysis is important; however, it isnecessary to evaluate the survivability characteristics of aircraft as well. Kim et al. [8]recently developed an algorithm to predict the susceptibility of combat aircraft usinga weighted score algorithm to assess the importance of various design parameters.

The IR signature can be predicted by analyzing the distribution of therear fuselage temperature. Numerical analysis of nozzle flow is a powerfultechnique employed to estimate the skin temperature. In this study, a mathematicalformulation and corresponding numerical schemes are introduced to evaluatethermal radiation shields, and parametric studies are performed whereby the heatshielding material characteristics are varied, as well as the dimensions of the shieldin order to predict the aircraft nozzle flow and the temperature distribution at therear of the fuselage.

2. NOZZLE FLOW ANALYSIS

2.1. Nozzle Flow

The engine layout studied in this paper is illustrated in Figure 1. It consistedof a jet nozzle, radiation shield, and the engine casing. A convergent–divergentnozzle shape [9] was considered for the supersonic flow, and the outermost duct ofthe engine casing was exposed to the freestream. The radiation shield and enginecasing were both considered to be thin shrouds. The radiation shield was positionedbetween the nozzle wall and the outer casing. Thermal radiation can be effectivelyreduced by placing more than one thin metal shroud between surfaces [10]. This canbe applied to the rear of an aircraft fuselage to reduce the heat transfer from thenozzle to the exterior of the fuselage [5]. Hot combustion products (FLUID 1) flowthrough the nozzle, while the two cooling air flows (FLUID 2) pass through insidethe radiation shield and the casing. The freestream (FLUID 3) exists at the outerparts of the engine casing; this describes the atmospheric conditions and the flightspeed of the aircraft.

To simulate the flow fields at the rear of the fuselage, the computationaldomain was divided spatially into four blocks, as shown in Figure 1b. The freestreamdomain extended in the radial direction sufficiently far to ensure that the boundaryconditions did not affect the region of interest. Infrared emissions from an aircraftcan be influenced by the nozzle exhaust plume as well as interior gas flow in thenozzle. In this study, the emphasis was on the temperature distribution of therear fuselage, which is affected by the gas flow within the nozzle. Therefore, thesimulation domain was restricted to the interior of the fuselage and the freestream.

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92 K. J. YI ET AL.

Figure 1. Schematic diagram of the rear of the fuselage showing the radiation shield. (a) Engine layoutand (b) multi-block grid system.

The general molecular formula for Jet A-1, which is the most commonly usedaviation fuel for gas turbine engines, is C11H22. On the assumption of stoichiometricreaction, the combustion chemistry is as follows.

C11H22 + 16�5 O2 + 3�76N2� → 11CO2 + 11H2O+ 62�04N2 (1)

Therefore, the combustion products can be assumed to be composed of carbondioxide, water vapor, and nitrogen, with the following molar fraction: 13.1% CO2,13.1% H2O, and 73.8% N2. The temperature of the gas mixture increases as the flowpasses through the combustion region in the engine. In this work, the pressure andtemperature at the nozzle inlet were set to 3.41 atm and 2,000 K, respectively. Thecooling air and freestream flow were modeled as ambient air, i.e., 21% O2 and 79%N2. The cooling air flowed along both the inner and the outer sides of the shield witha Mach number of 0.3. The freestream Mach number, which represents the flightspeed of the aircraft, was assumed to be 2.05. These conditions are summarized asfollows.

FLUID 1 : p = 3�41 atm� T = 2,000KFLUID 2 : p = 1 atm� T = 300K�M = 0�3FLUID 3 : p = 1 atm� T = 300K�M = 2�05

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HEAT SHIELDING IN JET ENGINE EXHAUST SYSTEMS 93

2.2. Boundary Conditions

In order to predict the nozzle wall temperature accurately, the heat transferbetween the gas and the nozzle wall and thermal conduction within the solidwall must be analyzed simultaneously. At the wall surface, heat is transferred byconduction and radiation. The heat balance can be described using the followingequation.

qc�gas + qr�gas = qc�solid (2)

where the subscripts c and r denote conduction and radiation, respectively. The twoterms on the left-hand side of the above equation describe conductive and radiativeheat transfer by the gas. The right-hand side of equation gives the heat conductedby the solid.

As the radiation shield and outer casing were both considered to be thinshrouds, heat transfer by the inner and outer flow field should be balanced with thatat each wall. The boundary conditions at each wall were as follows.

Nozzle inside wall surface: qc�F1 + qr�F1 = qc�wall (3a)

Nozzle outside wall surface: qc�F2 + qr�F2 = qc�wall (3b)

Radiation shield: qc�F2_i + qr�F2_i = qc�F2_o + qr�F2_o (3c)

Outer casing: qc�F2 + qr�F2 = qc�F3 + qr�F3 (3d)

In these relationships, the subscripts F1, F2, and F3 correspond to FLUID 1,FLUID 2, and FLUID 3, respectively. FLUID 2 flows along both the inner andthe outer sides of the radiation shield. The cooling air flowing inside the shield isdenoted F2_i, and that passing along the outer side of the shield is denoted F2_o.

3. NUMERICAL SCHEME

Pre-conditioned Navier–Stokes equations were employed to analyze the nozzleflow field; these equations are applicable to flows at all Mach numbers. TheAUSM+-up scheme was used to calculate the inviscid flux. A finite-volume methodwas used to deal with the radiative heat transfer by employing a weighted sum ofgray gases model.

3.1. Pre-Conditioning Algorithm

Subsonic and supersonic flows coexist inside the nozzle, and convergence isdecreased due to the difference in compressibility. A pre-conditioning scheme wasapplied to solve both incompressible and compressible regions by including anartificial viscosity term, which was based on the difference between the acousticand convective speeds. The pre-conditioning matrix by Weiss and Smith [11] wasused. This method converts the Navier–Stokes equations with conservative variables �� �u� �v� �e� into a form with primitive variables p� u� v� T� in order to avoid

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94 K. J. YI ET AL.

the density singularity problem. The pre-conditioned two-dimensional axi-symmetricNavier–Stokes equations are expressed as follows.

��Qv

�t+ � E − Ev�

�x+ � F − Fv�

�y= �Hs (4a)

where

Qv =

puvTkYi

� E =

�u�u2 + p�uv�uh�uk�u�uYi

� F =

�v�v

�v2 + p�vh�vk�v�vYi

� Ev =

0�xx�xy

u�xx + v�xy − qx�kx�xqix

Fv =

0�xy�yy

u�xy + v�yy − qy�ky�yqiy

� Hs =

1r

�v�uv�v2

�vh�vk�v�vYi

+ 1

r

0�xy

�yy − ���u�xy + v�yy − qy

�ky�yqiy

(4b)

Here, Qv is the vector of primitive variables, E and F are the inviscid fluxes, Ev

and Fv are the viscous fluxes, and Hs is an axi-symmetric vector. The above equationscan be applied to two-dimensional problems with alpha of zero, and they can also beapplied to axi-symmetric cases with alpha of unity. � denotes the pre-conditioningmatrix, which is defined as follows.

� =

��u�v

�h− 1�k��Yi

0�0�u000

00��v000

�T�Tu�Tv

�Th+ �Cp

�Tk�T�TYi

0000�00

00000�0

000000�

(5a)

where

� = 1U 2

r

− 1a2

+ 1RT

� �T = ��

�T

∣∣∣∣p

(5b)

A reference velocity, Ur , is defined to solve the convergence problem instagnation regions [12]. By using the above pre-conditioning matrix, robustness andconvergence of the calculation can be guaranteed for flows at all Mach numbers.

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HEAT SHIELDING IN JET ENGINE EXHAUST SYSTEMS 95

3.2. AUSM+-Up Scheme

The AUSM+-up scheme is part of the modified advection upstream splittingmethod (AUSM) family of schemes, and was used to approximate the inviscid fluxesin the Navier–Stokes equations. AUSM schemes have the advantage of providing arapid numerical resolution of strong shocks and accurate results for boundary layers.This method can be applied not only to low Mach number flows, but also to multi-phase flows. The AUSM+-up scheme is a refinement of the AUSM+ scheme that wasdeveloped to be reliable over arbitrary fluid speed ranges [13]; it provides improvedrobustness for the problem described here.

The inviscid fluxes in the Navier–Stokes equations are explicitly split into twoparts, describing convection and pressure [14]. As such, it is sufficient to consideronly the one-dimensional flux, excluding turbulence and energy terms, termed E′

here, to briefly introduce the fundamentals of AUSM+-up scheme. The inviscidflux at the interface of the control volume, denoted by subscript “1/2”, is given asfollows.

E′1/2 = m1/2L/R + P1/2 (6a)

where

= 1� u�H�T � P = 0� p� 0�T � m1/2 = u1/2�1/2 = a1/2M1/2�L/R (6b)

♦�L/R ={ ♦�L if u1/2 > 0 ♦�R otherwise

(6c)

In the above expressions, the flow conditions L and R at the cell interfaces can bedetermined using the Chakravarthy method with the minmod slope limiter [15].

In the AUSM+-up scheme, we set the interface Mach number and pressure interms of M1/2 and P1/2 to be as follows.

M1/2 = �+ 4� ML�+�−

4� MR�+Mp (7a)

P1/2 = �+ 5� ML� PL +�−

5� MR� PR + Pu (7b)

where the split Mach numbers with fourth degree polynomials and the split pressureswith fifth degree polynomials are defined as follows.

�± 4� M� =

{�±

1�� if �M� ≥ 1

�± 2�

(1∓ 16��∓

2�

)�− 1

16 ≤ � ≤ 12 � otherwise

(7c)

�± 5� M� =

{ 1M�±

1�� if �M� ≥ 1

�± 2�

[ ±2−M�∓ 16�M�∓

2�

]�− 3

4 ≤ � ≤ 316 � otherwise

(7d)

with

�± 1� M� = 1

2 M ± �M�� ��±

2� M� = ±14 M ± 1�2 (7e)

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96 K. J. YI ET AL.

Here, the pressure diffusion term, Mp, and the velocity diffusion term, Pu, areintroduced to improve calculations of low Mach number or multi-phase flow.

3.3. Methods for Considering the Radiative Transfer Equation

Radiation is a key element in the heat transfer of hot gas flow, and must beconsidered. The radiative transfer equation (RTE) governing the radiation intensity,I , for a medium at any position �r along a path �s through an absorbing and emittingmedium is given by the following.

dI �r� �s�ds

= −�aI(�r� �s)+ �aIb �r� (8)

The medium was assumed to be non-scattering. Scattering characteristics can beignored for an aircraft engine system, since the emission of soot particles thatscatter the radiant energy is not important. In the above equation, �a is theabsorption coefficient. A real gas is a non-gray gas with radiative properties wherethe absorption and scattering coefficients vary across the electromagnetic spectrum.Smith et al. [16] developed a set of absorption coefficients for the weighted sum ofgray gases model (WSGGM), in which a non-gray gas is replaced by a number ofgray gases. In this work, we used an absorption coefficient model assuming a mixtureof gases using a WSGGM with four gray gases.

The finite-volume method (FVM) is one of the most commonly used methodsfor solving the RTE [17–19]. In the FVM, we assumed that the magnitude ofintensity was constant, but the direction varied within the control volume andcontrol angle [20, 21]. Our discretization scheme was as follows:

amnP Imn

P = ∑I=E�W�N�S

amnI Imn

I + bmnP (9a)

where

amnI = −�AiD

mnci�in (9b)

amnP = ∑

i=e�w�s�n

�AiDmnci�out + �a�P�V��

mn (9c)

bmnP = �aIb�P �V��

mn (9d)

Dmnci =

∫ n+

n−

∫ �m+

�m−

(�s · �ni

)sin�d�d (9e)

��mn =∫ n+

n−

∫ �m+

�m−sin�d�d (9f)

Here, �Ai, �V , and ��mn are the surface area, control volume, and control angle,respectively. The directional weight, Dmn

ci , determines the inflow or outflow of radiantenergy across the control volume face.

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HEAT SHIELDING IN JET ENGINE EXHAUST SYSTEMS 97

4. RESULTS AND DISCUSSION

4.1. Code Validation

We developed simulation software based on the methods described above. Toestablish the reliability of these formulations, a numerical analysis of the nozzledesigned by the Jet Propulsion Laboratory (JPL), which is widely used for validationof nozzle flow analysis calculations, was performed. We compared our results withexisting simulated [22] and experimental [23] data. The working fluid inside thenozzle was assumed to be air, and the initial pressure and temperature were set toP0 = 1�0342MPa and T0 = 555K, respectively. The pressure distributions along thewall and centerline are shown in Figure 2, which indicated good agreement withexisting data for the same system.

4.2. Effect of the Radiation Shield

The initial conditions of the combustion products, cooling air, and freestreamare listed in section 2.1. The heat conduction phenomena at the nozzle wall wasmodeled assuming a thermal conductivity of 20 W/(mK), while the shield and outercasing were assumed to be thin shrouds; thus, the material characteristics of thesesections were not considered.

Before checking the effect of the radiation shield on the skin temperaturereduction, we have looked into the flow phenomena. Figure 3 shows the Machnumber and pressure distributions. As the gas flowed through the convergent–divergent nozzle, the flow was accelerated and became supersonic. As illustrated inFigure 3, at the throat, the flow near the nozzle wall had a higher Mach number

Figure 2. Comparison of the pressure variation along the wall and centerline for the well-studied JPLnozzle flow.

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98 K. J. YI ET AL.

Figure 3. Contour plots of the (a) Mach number and (b) pressure.

than near the centerline. This was because the expansion of gases passing throughthe throat region occurred more rapidly near the wall than at the nozzle axis. Thisphenomenon was also observed in the JPL nozzle flow shown in Figure 2. In thefreestream domain, an expansion wave was created along the nozzle wall, and theMach number increased as the pressure decreased.

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HEAT SHIELDING IN JET ENGINE EXHAUST SYSTEMS 99

Figure 4 shows the heat flux variations without shield in Figure 4a and withshield in Figures 4b and 4c. As mentioned earlier, the radiative heat flux is observedto be predominant over the conductive heat flux to casing as well as shield. Themaximum radiative heat flux is almost 2.7 times greater than that of conductiveheat flux at the inner wall of shield which is affected by hot combustion productsinside the nozzle, as shown in Figure 4b. In comparison of Figure 4a with Figure 4c,by placing the shield between the nozzle wall and the outer casing, the radiative aswell as conductive heat flux from inside of engine to the outer casing can be greatlyreduced.

The temperature distribution along each component of the rear of the fuselageis shown in Figure 5. With no radiation shield, which is depicted by the solid line,as the hot combustion gas entered the nozzle inlet, heat was transferred outwardand the temperature of the outer wall of the nozzle and the outer casing increased.When the flow accelerated inside the nozzle, the pressure and temperature dropped,and therefore the outer casing temperature decreased along the direction of the flow.The cooling air inside the casing also affected this temperature drop. At the innerwall of the nozzle (unlike the nozzle outside wall and casing), the evolution of thetemperature profile changed between the start of the converging section and thethroat. This occurred irrespective of the existence of the shield because of the changeof the nozzle wall thickness. Since the nozzle wall acted like a thermal resistance,as the thickness of nozzle wall increased, the amount of heat escaping was reducedand the temperature of the inner wall rose. However, there was a point where thetemperature decreased suddenly at the location where the nozzle throat began tocurve. Here, the pressure near the wall fell rapidly and the temperature also dropped.

Now, we consider the temperature profile with and without the radiationshield. As shown in Figure 5, even if the outer casing was placed at the samelocation, the temperature profile depended on the domain of the radiation shield.For the case without the shield (shown by the solid line in the figure), the casing hada temperature range of 550 K to 950 K, whereas with the shield (shown by the dash-dotted line in the figure), the temperature ranged from 370 K to 640 K. By placingthe radiation shield between the nozzle wall and the outer casing, the temperature atthe rear of the fuselage was reduced by 26 to 40 percent.

4.3. Effect of Material Characteristics

In this section, the effect of the thickness of the radiation shield on thetemperature of the rear of the fuselage is examined. The thickness of the radiationshield and casing should be considered, as well as that of the nozzle wall. Thethickness of the shield and casing were set to be same as that of nozzle wall inlet, asillustrated in Figure 6, and the other conditions were kept the same as described insection 4.2.

In the previous section, we assumed that the radiation shield and enginecasing were thin shrouds, and only the thermal conductivity of the nozzle wall wasconsidered. However, here we consider thermal conduction in all three walls, andset the thermal conductivity of each to be 20 W/(mK), independent of temperature.Figure 7 shows the temperature distributions of each component when all the wallshad a finite thickness. Compared with Figure 5, the tendency of the temperature

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100 K. J. YI ET AL.

Figure 4. Heat flux variations (a) without shield and (b) and (c) with shield.

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HEAT SHIELDING IN JET ENGINE EXHAUST SYSTEMS 101

Figure 5. Effect of the radiation shield on the skin temperature.

profiles of each wall was the same, regardless of the thickness. In this case, boththe radiation shield and the casing had different inner and outer walls temperatures.For the casing wall, the temperature difference between the inner and outer wall wassmaller than for the shield, because the temperature difference across this componentwas smaller. It follows that the material characteristics of the radiation shield weremore important than those of the outer casing.

Now, we examine the effect of the thermal conductivity of the radiationshield on the skin temperature. The thermal conductivity of the nozzle wall andcasing were kept at 20 W/(mK); however we varied the thermal conductivity of the

Figure 6. Schematic diagram of the computational domain when the radiation shield and the outercasing have a finite thickness.

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Figure 7. Effect of the thickness of the radiation shield and the outer casing on the skin temperature.

radiation shield. The materials used for an aircraft fuselage need to be strong, hard,and lightweight; duralumin and titanium are typically used for these applications.Duralumin is an alloy of aluminum with copper, magnesium, and manganese, and iscommonly used in the aircraft industry because of its excellent corrosion resistanceand high specific strength. Titanium has a low thermal conductivity and excellentspecific strength. Figure 8 shows the temperature profiles when the radiation shieldwas made of titanium or duralumin. The thermal conductivity of titanium andduralumin are 20 W/(mK) and 141 W/(mK), respectively. As the conductivityincreased, more heat was transferred through the heat shield; thus, the temperatureof the inside wall became lower and that of the outside wall became higher. Asshown in Figure 8a, for the duralumin case, the shield transferred more heat, andthe outside wall temperature of the shield increased compared to the titanium case.Because the temperature of the outer casing was affected by the shield temperature,the casing temperature with the duralumin shield was higher than that with thetitanium shield, as seen in Figure 8b. To reduce the temperature of the outer duct, alow-conductivity metal should be chosen for the radiation shield.

4.4. Effect of the Radiation Shield Length

In order to investigate the relationship between the length of the radiationshield and the temperature of the rear of the fuselage, we altered the computationaldomain of the radiation shield, as shown in Figure 9. The engine layout with aradiation shield extending to the nozzle exit was termed A-type, while that witha shield that reaches the starting point of the curvature of the nozzle outer wallwas termed B-type. For both engine layouts, the shield and the outer casing wereconsidered to be thin shrouds to examine only the effect of the shield shape on thetemperature distribution.

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Figure 8. Effect of the thermal conductivity of the radiation shield on the (a) shield temperature and(b) casing temperature.

Figure 10 shows the temperature distributions obtained under the same flowconditions described in the previous sections. For the B-type shield, the temperatureof the outermost duct increased rapidly at the end point of the shield. This wasbecause the radiation shield blocked heat primarily from the nozzle ends along theflow; thus, heat was transferred directly to the outer casing downstream from the

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Figure 9. Illustration of the different lengths of the radiation shield.

point where the heat shield terminated. The decrease in the temperature at the tipof the shield can be explained by considering that the two cooling air flows – oneinside the shield and one inside the casing – mix at the shield tip, and this mixingresults in a reduction of the skin temperature.

As well as presenting a comparison between A-type and B-type heat shields,we also compared these engine layouts to one with no radiation shield, composedof a nozzle wall and an outer casing only. The outermost duct was placed at samelocation as in the previous case. The resulting temperature distribution is illustratedby the long-dashed line with circle symbols in Figure 10. This casing had a highertemperature in the upstream region, which decreased along the direction of the gasflow. In the B-type design, the shield was placed partially over the hot region; inother words, the shield was expected to reduce the skin temperature most effectivelyby placing it close to regions of high temperature. Consequently, the temperature at

Figure 10. Effect of the length of the radiation shield on the skin temperature.

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HEAT SHIELDING IN JET ENGINE EXHAUST SYSTEMS 105

the rear of the fuselage could be reduced effectively with minimal drawbacks causedby the radiation shield (i.e., weight increase and thrust loss).

5. CONCLUSION

We examined the effects of a heat shield on the survivability of an aircraftby performing a numerical analysis of the gas flow at the nozzle of turbofan jetengines. Reducing the temperature at the rear of the fuselage is a critical factor inimproving aircraft survivability. The numerical simulations focused on the effects ofthe nozzle wall, radiation shield, and casing, and on the thermal transport propertiesof the engine in order to determine the skin temperature of the rear of the fuselage.Hot combustion products pass through the nozzle and transfer heat to the surfaceof the nozzle wall as well as the ducts covering the nozzle. The temperature of theoutermost duct was affected by the material properties of the heat shield and by theengine layout, leading to a change in the skin temperature observed from outside.

To reduce susceptibility, we investigated the effect of the radiation shieldplaced between the nozzle wall and the casing. The material characteristics of therear of the fuselage affected the wall temperature; a material with low thermalconductivity transferred less heat through the wall, and thus the outside walltemperature was reduced and the temperature of the outer casing was lower.

We also studied the effect of the spatial extent of the heat shield withinthe engine. The point to which the shield extended influenced the heat transferphenomenon inside the engine system, but also it added mass to the aircraft. Byplacing the shield only over the hottest regions of the exhaust system, the skintemperature could be decreased effectively while minimizing the additional aircraftweight caused by the heat shield.

FUNDING

This work has been supported by the Low Observable Technology ResearchCenter program of the Defense Acquisition Program Administration and Agency forDefense Development.

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