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Design and Wind Tunnel Study of a Top-mounted Diverterless Inlet
V ol. 17
No. 2
CH INES E JO U RN A L O F AER ON A U TICS
M ay 2004
Design and Wind Tunnel Study of a Top mounted Diverterless Inlet T AN Hui jun, GU O Rong wei ( I nternal Fl ow Research Cent er , Nanj ing Uni versity of A eronautics and A str onautics , Nanj i ng Abstract:
210016 , China)
Combined w ith a U A V o f the shape like Global Hawk, a new inlet is advanced to obtain
hig h performance in both Radar Cross Section( R CS) and aerodynamic drag. Efforts are made to achieve t his goal such as ado pting a top mounted inlet configuration, ut ilizing the diverterless technique and putting forw ar d a new shape of entrance. A design method is brought for ward and v er ified by wind tun nel tests. Results indicate: ( 1) Despite the negativ e effect of the fr ont fuselage and the absence of the conventional boundar y diverter , t he performance of the top mounted diverterless inlet advanced here ( M a: 0 50 0 70,
:- 4 6,
> 0 975) is equivalent to that o f convent ional S shaped inlet with di
verter; ( 2) T he integration of the inlet w ith the fuselage is realized by t he utilization of a special inlet section and t he div erterless technique, w hich disposes t he whole inlet in the shield of the head of UA V, improving the drag char acteristics and the stealthy performance of the aircraft; ( 3) T he bump which is equal to the local boundary layer thickness in height can diver t the boundary layer effect ively. As a re sult, no obv ious low total pressure zone is found at the out let of the inlet; ( 4) Accor ding to the ex peri mental results, negative angle of attack is favorable to the total pressure recov er y and positive ang le of at tack is favorable to the total pressur e distortion, while y aw brings bad effects on bo th; ( 5) T he design of cowl lip is of great impo rtance to the inlet performance at y aw , therefore, further impro vement of the inlet performance will rely on the lip shapes of the cowl chosen. Key words:
top mounted inlet; div erterless inlet; unmanned air vehicle; design; w ind tunnel test
一种无人 机用背负式无附面层隔道进气道的设计及实验验证. 谭慧俊, 郭荣伟. 中国航空 学报( 英 文版) , 2004, 17( 2) : 72- 78. 摘 要: 结合一类似于 全球 鹰! 的无 人侦察 机外形, 对一 种新型 高隐身 低外阻 进气道进 行了如 下 设计: 采用背负式布局方案, 使用无隔道技术, 并提出了一 种新的进口 截面形状。加 工了风洞 试验 模型并开展了验证性风洞实验研究工作。结果表明: ( 1) 尽管受到 机头遮蔽 的不利影响, 且没 有采 用传统的附面层隔道, 所给出的背负式无隔道 进气道方 案性能( Ma: 0 50~ 0 70,
:- 4 ~ 6,
>
0 975) 与常规的有隔道 S! 弯进气道相当; ( 2) 特殊进口截面形 状及无附 面层隔道技 术的采用 将进 气道与机身有机地融为一体, 使进气 道整体 都处于 飞行器头 部的遮 蔽之中, 这 有利于改 善飞行 器 的阻力特 性和隐身性能; ( 3) 进气道出口截面上未发现因附面层吸 入而造成 的低总压区, 这说 明高 度与当地附面层厚度相当的进口鼓包能有效地隔除附面层中能量较低的气流; ( 4) 研究范围内, 负 攻角对进 气道的总压恢复系数有利, 正攻角 对周向 畸变指数 有利, 而侧 滑角则 对两者均 有着不 利 影响; ( 5) 唇口的设 计对进气道的侧滑角性能有着重要影响, 进气道性能 的进一步提高应考虑唇口 设计的改进。 关键词: 背负式进气道; 无隔道进气道; 进气道; 无人机; 设计; 实验 文章编号: 1000 9361( 2004) 02 0072 07
中图分类号: V211. 3
文献标识码: A
T op mounted inlets have advant ages over con vent ional inlet conf igurat ions as leav ing more space
creasing st ealt hy capabilities t o up looking radar, and prevent ing foreign objects from ing est ing int o
underneat h t o place missiles or ot her weapons, pro viding bett er environments for g round crew , in
eng ines[ 1] . T hey are quit e common on commercial aircraf ts w hen t hree engines are desired like on t he
R eceived dat e: 2003 05 19; R evision received dat e: 2004 02 10
M ay 2004
D esign and W ind Tunnel St udy of a T op M ount ed Divert erless Inlet
# 73 #
M cDonald Doug las DC 10, MD 11. However, t he complex aerodynamics issues, especially at high an g les of att ack, impeding t he top mounted inlets∀ w ide use on milit ary aircraft s, and at present t hey are mainly applied t o UCAVs and bombers such as
Fig. 1
Sketch of the fuselag e
Global Hawk, X 45, X 47 and B 2. In t he fift ies of last century , the explorat ions of the divert erless t echnique w ere init iat ed by re
2
Design of t he Inlet
searchers in NACA Lew is Laborat ory [ 2] . But t he
T he bulge at the front fuselage shadows t he
diverterless technique did not get furt her at tent ions
top mounted inlet and is believed to reduce t he
unt il its recent successful applicat ion in X 35. T his unique technique, eliminat ing t he divert er and ot h
front al aspect Radar Cross Sect ion ( RCS ) of t he
er f eat ures associat ed w ith boundary lay er m anage
inlet . H ow ever, it is a great challenge t o design t he inlet because t he bulge speeds up the development
ment, is realized by desig ning a three dimensional
of t he boundary layer, resulting in the apert ure of
surface, or bump, f ront of the inlet apert ure, w hich can decelerat e part of t he com ing f low and
the inlet mostly in t he boundary layer. In order t o
create a pressure gradient that pushes the boundary
solve the problem mentioned, t he Global Haw k adopt s hig her boundary diverter, w hich lift s t he
layer aw ay from the inlet . T hus, t his technique is
inlet to ameliorate the f low condit ion in front of t he
able to reduce t he overall w eight of t he aircraf t, cut dow n t he front al area of the aircraf t[ 3] , improve
aperture. But t his approach w ill increase RCS and the aerodynamic drag of t he aircraf t. Hence, a top
the drag characterist ics and enhance t he st ealt hy
mount ed inlet int egrated w it h the f uselage is
capabilit ies by decreasing corner reflections of radar
brought forward to dispose t he w hole inlet in t he
w aves. According to dom est ic open literat ure, lit t le
shield of t he bulge. T he diverterless t echnique is u t ilized in this design t o deal w it h the t hick bound
w ork has been made on t he top mount ed inlet or
ary flow and t o obt ain bet ter conformance. Addi
the divert erless technique, let alone t he combina
t ional endeavours are also m ade to acquire f avorable
t ion of both. Integ rated w ith a UAV of t he conf ig urat ion like Global Hawk, a top mount ed S shaped
inlet perf orm ance. T hough the conventional design method[ 4] of
inlet utilizing the diverterless t echnique is dev ised
S shaped inlet is still used here, a creat ive cross
and verif ied by w ind tunnel t ests in t his study.
sect ion is come up w ith, w hich conveniently inte
Description of the Forebody
g rate t he inlet w it h t he f uselage and the bump em ployed by t he divert erless t echnique. T he proce
F ig . 1 demonst rat es t he f orepart of an aircraf t,
dure of t he inlet design is described in det ail in t he
1
w hich resem bles t he Global Haw k, the UAV of
follow ing.
U S Air Force. T he dist ance from t he apex of t he fuselage to t he leading edg e of t he inlet is L . As
2. 1
shown in t he f ig ure, a bulge which is 0 06L in
as lengt h, offset and exit area are subject t o geo
height rests on the top side of the fuselage. Because
metric const raints of t he fuselage and t he engine,
the w ing s are below and far aw ay f rom the inlet , it is obvious t hat t hey have litt le inf luence on the per
they are alw ays know n w hen one set out to design a inlet. According to a selected area rat io, t he en
f orm ance of t he inlet at low incidence. Accordingly
t rance area of t he inlet can t hen be decided. T he
the design and w ind t unnel test s of the inlet in t his paper have not taken the eff ect of the wings int o
paramet ers used in t his paper are summarized in T able 1, t wo of w hich are normalized by t he diam
account .
eter of the ex it .
General geometrical parameters of the inlet F or some general parameters of the inlet such
# 74 #
T A N Hui jun, GU O Rong w ei
Table 1 General geometrical parameters of the inlet
2. 4
2. 2
1 6
1 3
1 5
Design of the entrance shape T he reason w hy t he convent ional cross sec
t ions like ellipse or racew ay shape can not be used
Design of the diffuser T he cross sectional area ( A i ) at arbit rary sta
Lengt h/ Diameter O f fset / Diameter A rea rat io Equivalent angle/ ( ) 5 0
t ion of the diff user can be obt ained by its entrance area, exit area and area dist ribution. Ref erring t o Fig. 2, it is not diff icult t o deduce t he mat hemat i cal expression of any cross sectional area
here is t hat the diverterless technique requires t he
Ai =
integrat ion of the entrance section w it h t he bump, w hile neither of t hese can sat isfy this requirement . T he new ly dev ised cross sect ion is pict ured in F ig. 2, w hich consist s of a cirque ( centre ang le , in t ernal radius r and ext ernal radius r + a + b) and four quarter ellipses ( major axis a or b and minor
i
2
( 1)
ai bi ri ,k = ,k = and assum ing ci 2, i c i 3, i ci these rat ios and i are know n, Eq. ( 1) can be rew rit ten as Lett ing k 1, i =
ci =
its, one is t hat it can be int egrated w ith the bump and t ransformed to the shape of t he engine f ace
+
easily, and t he ot her is that it unit es t he inlet w ith aerodynam ic drag and low observ abilit y.
!a ici !bic i i + + !( r i + a i + b i ) 2 2 2 360 - 360!r i
ax is c) . For t his configuration t here are t wo mer
the fuselage t o the ut most deg ree achiev ing low
CJA
i
360
Ai M
M= !
k l , i + k 2, 2
( k l , i + k 2, i + k 3, i ) 2 -
i
360
i
k 23, i
which means t hat the configuration of this sect ion is ex act ly determined. k 1, i , k 2, i , k 3, i and i vary smoot hly and monotonously from k 1, in , k 2, in , k 3, in and
in
to k 1, out, k 2, out, k 3, out and
out
when t he sta
t ion moves from the ent rance to the ex it . T he val ues of k 1, in , k 2, in , k 3, in and in det erm ine the shape of t he entrance section, which are select ed 1 0, 2 5, 2 5 and 110 F ig. 2 Sketch of the entrance section
2 3
Centerl ine and area distributions T he centerline and t he area distribut ions are of
g reat significance to the performance of the inlet. Centerline shapes determ ine t he turning of the in t ernal f low and t hereby decide t he transverse pres
individually in this st udy.
While determining k 1, out, k 2, out , k 3, out and out one should be more careful to guarant tee t hat the exit of t he dif fuser is an ex act circle. T he values of 1 0, 1 0, 4 0 and 0 are capable to meet the re quirement well and are used here. Result show s that t he boundary condit ions of k 1, i , k 2, i , k 3, i and i
work w ell ( F ig . 3) .
sure gradient and t he secondary flow in t he duct. Area dist ributions det erm ine the diff usivit y and hence control t he streamw ise pressure gradient im posed on the flow. Inappropriate cent erline shape or area dist ribut ion may both lead t o f low separa t ion. Ref. [ 4] provides three poly nomial funct ions w hich are suit able for both centerline shape and area distribut ion. Based upon previous investiga t ions, the cent erline distribut ion w it h a modest
F ig. 3 3 D view of the diffuser
2. 5
Design of the bump
turning and t he area dist ribution w it h a rapid
T he success of t he diverterless technique w ill be largely dependent on t he design of t he bum p. A
change near t he ex it are used in this study.
well designed bump should ex ert its f unct ions of
M ay 2004
D esign and W ind Tunnel St udy of a T op M ount ed Divert erless Inlet
# 75 #
both guiding and pushing to keep the boundary lay
the innovat ive cross sect ion of the inlet , an int egra
er aw ay f rom t he inlet by elaborat e surf ace design.
t ion of the inlet and t he fuselage is obt ained w hich
T he effect of guiding depends directly on t he bump it self and operat es on t he low er part of t he bound
minimizes t he inlet∀ s heig ht and t hereby brings t he whole inlet into the shadow of t he bulg e of t he
ary layer flow . But the ef fect of pushing appears
fuselage.
not so direct and is det ermined by the pressure gra dient formed by t he bump act ing on t he upper lay er . Wit h a crescent shaped cross sect ion and a polynomial vert ical section line, t he bump is ap prox imate t o a cone shaped surf ace, w hich is t an gent to t he fuselage and t he inlet at it s boundaries. Computational result s indicate t hat bet ter perf or mance of the inlet can be achieved by a bump of F ig. 5 Enlarg ed view of the inlet∀ s forepart
w hich t he max imum heig ht is equivalent to t he thickness of the local boundary layer ( evaluat ed by
In addit ion, t he junct ion of t he inlet and t he
turbulent boundary layer of flat plate ) and t he
eng ine ( w it h center body) is designed using con
lengt h is about five times the max imum height . 2. 6 Design of cowl lips
st ant area dist ribut ion. At the same t ime, t he de sign of outside surface of the inlet is also conduct
In general, lips of subsonic inlets are of impor t ance t o the performance of t he inlet. And, w it hal,
ed.
3
in t his paper it act s as an addit ional measure assist ing the bump t o push boundary f low aw ay from t he inlet. A lip scheme wit h variable section line and [ 5]
v ariable lengt h is adopt ed. NACA 1 85 100 lip
is
Experimental Verification
In order to obtain t he aerodynamic perf or mance of t he new ly designed inlet and verify t he design approach, an experiment al inlet model is de
at top point of t he ent rance ( T point in F ig . 2) forming a local ∃dif fuser∀ ( F ig. 4) to acquire high
signed and manufactured and t hen test ed in a wind tunnel.
er pressure in reg ion A t han t hat of t he outside. Sym met ry airfoil is at t he intersect ion point w ith
3. 1
the bump ( S point in F ig . 2) to prevent perf or mance deterioration at yaw . T he ut ilization of vari able lengt h lip gets a forw ard sw ept cow l which co
Inlet model and test tunnel
T he complet e t est model consist s of fuselage, top mounted inlet, eng ine cent erbody, pressure rakes and conic flow plug as illustrated in Fig. 6.
operat es w ith the local ∃ diffuser∀ and ensures t he
Due to the constraints of w ind t unnel blockage and the m easurement of internal f low , t he w ings are
majorit y of the boundary layer f low spilling out at the aft notch.
not taken into considerat ion. T he t est model has a tot al length of 1000mm and the diameter at t he outlet of the inlet, where the measurement of t he pressures is made, is 40m m.
Fig . 4 F eature o f the upper lip and the inlet bump
F ig. 6 Sketch of the test model
F ig. 5 gives an enlarg ed view of t he inlet ∀ s forepart. In virtue of t he divert erless technique and
An eight leg rake is inst alled at t he out let of the inlet for measurement of st eady tot al pressure.
# 76 #
T A N Hui jun, GU O Rong w ei
CJA
Each leg has five probes, providing a tot al of 40 lo
wards. T he possible reason may be that at low
cat ions. Four st at ic pressure taps are locat ed on t he
mass f low rat io more low energy flow near t he
inner surface of the inlet equally along the circum
fuselage enters t he inlet relat ive t o t he total mass
f erence and in a plane cont aining the face of the t o t al pressure probes. T he m ass f low plug , placed at
flow captured. ( 2) Ef fects of free st ream Mach number
the model ex it and driven by an electric mot or,
F ig. 8 shows the variat ion of tot al pressure re
prov ides flow cont rol for t he measurement s at dif f erent mass flow rat ios.
covery coeff icient w it h f ree st ream M ach number. It can be noted t hat the recovery coef ficient is in
T he experiment al study is accomplished at
sensitive t o M ach number up to 0 7 and remains
NH 1 high speed wind t unnel of NUAA. T he
above 0 975 st ably. T his indicat es that , despite
Reynolds number of t he w ind t unnel ranges f rom 1 2 % 107 to 1 77 % 10 7 . During this t est, t he free
the neg at ive impact of t he bulge and t he absence of the boundary diverter, the recovery coeff icient of
st ream Mach number varies f rom 0 5 t o 0 8, ang le
the inlet tested is equivalent t o that of convent ional
of at tack from - 4 t o 6 and yaw from 0 to 2 . T he maximum blockage of the t est sect ion is 4 3%
S shaped inlet w it h diverter. How ever, w hile t he Mach number is greater than 0 7, t he recovery co
w hile t he inlet model is inst alled.
eff icient t akes on a sharp decline. T his t rend is
3 2
probably aroused by a local shock occurring at t he
Experimental resul ts
T he performance of the inlet is evaluat ed by the t otal pressure recovery coef ficient ( ) and cir
back of t he bulge, w hich thickens rapidly t he boundary layer and det eriorat es t he flow condit ion
cular t otal pressure distort ion ( ∀ ) . In the f ollow
front of the inlet. It also means that t he fuselage
ing text , t he eff ects of mass flow ratio, f ree st ream M ach number, angle of at tack and yaw are illus
conf igurat ion like Global H aw k is not suitable for high subsonic flight .
t rat ed. Cont ours of tot al pressure at the outlet of the inlet are also presented. ( 1) T ot al pressure recovery coef ficient at dif f erent mass flow rat ios In Fig. 7, t he tot al pressure recovery coeff i cient is plott ed as a funct ion of mass f low rat io at 0 ang le of att ack and yaw . It can be seen t hat , at first , w it h t he descent of mass f low rat io t he recov ery coeff icient ascends due t o the reduction of t he frict ion loss. But when the m ass flow rat io goes dow n further the recovery coeff icient t ends dow n
Fig . 7 T o tal pressure r ecovery coefficient v ersus mass flo w ratio( = 0 , #= 0 )
Fig . 8 T o tal pressure recovery coefficient versus free stream M ach number
Fig. 9
Cir cular distortion coefficient versus free stream Mach number
M ay 2004
D esign and W ind Tunnel St udy of a T op M ount ed Divert erless Inlet
# 77 #
For all test M ach numbers, t he circular total
trance of the inlet is lower t han that of free st ream
pressure distortion keeps below 1. 2% , as show n in
because of the influence of the upper fuselage,
F ig. 9. It means t hat the dist ribut ion of tot al pres sure losses is nearly constant along circular direc
which results in a f ew effect on circular distort ion. For t he case of posit ive angle, because the count er
t ion.
rot at ing vort ex pair developed by t he up w ash f low around t he fore fuselage sweeps off part of t he low
( 3) Eff ect s of incidence T he model is tested at six ang les of at tack, of w hich t he range is f rom - 4 t o 6 . Fig. 10 and F ig. 11 present t he ef fects of ang le of at tack on t o t al pressure recovery coef icient and circular total pressure dist ortion respectively. It can be observed from t hese t wo f igures that negative angle of at tack benefit s recovery coef ficient and aff ect s distort ion slightly, w hile posit ive angle of at tack affects re covery coefficient slight ly and improves distort ion obv iously. For t he case of negative ang le, it can be
energy flow in front of t he inlet , t he curve of t he tot al pressure recovery goes up a bit f rom 0 t o higher incidence and hence a remarkable drop of circular dist ortion occurs. ( 4) Ef fects of yaw T he ef fects of yaw on t otal pressure coeff icient and circular tot al pressure distort ion are illustrated in F ig. 12 and Fig. 13. It can be seen that 2 y aw has lit t le ef fect on recovery coeff icient but increases
explained as follows: on t he one hand, w it h nega
dist ortion distinctively ( still less t han 2% ) . More over, w ith t he increase of t he Mach number, t he
t ive angle of att ack rising, the ent rance of t he inlet is exposed t o f ree stream more and more, result ing
performance of the inlet is det eriorating , in part ic ular, at hig h M ach number, which can be associat
in more main flow captured and thereby t he in crease of tot al pressure recovery coef ficient; on t he
ed w it h unfavorable f low occurring near the lip at
ot her hand, the local flow ang le of at t ack at the en
higher M ach number. Consequently, t he improve ment of perf orm ance at y aw w ill lies on the melio ration of cow l lips.
F ig. 10
T otal pressure recovery coefficient versus angle of attack
Fig . 12 Effects of yaw on to tal pressure recovery co efficient
Fig . 11 Circular total pressure distortion co efficient versus ang le of attack
F ig. 13 Effects of yaw on circular distortion coefficient
# 78 #
T A N Hui jun, GU O Rong w ei
CJA
( 5) T otal pressure contours
technique are reported. During the design process,
F ig . 14 demonst rat es the t otal pressure dist ri
the int eg rat ion of the inlet w it h t he fuselage is real
butions at t he outlet of t he inlet w hile Ma = 0 7,
ized by t he utilizat ion of a special inlet sect ion and
= 0 , #= 0 . T he t ot al pressure cont ours make a feature of high m agnit ude and uniform distribut ion
elim inat ion of convent ional diverter, w hich dispos es t he whole inlet in the shadow of the bulge im
and no obvious low tot al pressure reg ion ex ists. All
proving t he drag characterist ics and t he st ealt hy
of t hese suggest t hat t he diverterless t echnique used in this paper can ef fectively push the low energ y
performance of t he aircraft . Wind tunnel tests indi cat e t hat the perf orm ance of t he inlet advanced here is equivalent to t hat of convent ional S shaped inlet
flow aw ay from t he inlet , w hich is validated by CFD result s ( F ig . 15) [ 6] .
wit h diverter, so t he diverterless t echnique adopted is successful.
References [ 1]
V ooren A , Kewenter H, Edefur H . T op mount ed air inlet for supersonic & transonic milit ary aircraft [ OL] . ( 2001 05 16 ) ht t p: / / w w w . ave. kt h. se/ education/ msc/ cours es/ 4E1232/ 2001/ aero2/ T M I R eport . pdf.
[ 2]
S imon P C, Brow n D W, Huf f R G . Performance of ext ernal compression bump inlet at M ach numbers of 1 5 t o 2 0 [ R ] . N A CA RM E56L19, 1957.
[ 3]
M cFarlan J D. Joint strike f ighter diverterless supersonic inlet [ O L ] , ( 2000 11 15) . ht tp: / / w w w. codeon emagzine. com/
F ig. 14 Contours of total pr essure recov er y coefficient
archieves/ 2000/ art icles/ july 00/ divertless
at the outlet [ 4]
1. ht ml.
Lee C C, Boekicker C. Subsonic dif fuser design and perf or man ce f or advan ced fight er aircraft [ R ] , A IA A 85 3073, 1985.
[ 5]
R ichard J R . A n invest igation of several NA CA series inlets at M ach number from 0 4 to 1 29 for mass f low rat io near 1 0 [ R ] . N AS A TM X 3324, 1975.
[ 6]
谭慧俊, 郭荣伟. 背负式无 隔道进气 道的方案 设计与性 能 研究[ R ] . 南京航空航天大学科技报告, R 0202018, 2002. T an H J, Guo R W. Design an d performance st udy of a top mounted diverterl ess inlet [ R ] . T echnical Report of N U AA , R 0202018, 2002. ( in Ch inese)
Biography: Fig . 15
Streamlines released from points 2mm away from fuselage
4
Conclusions
Design and w ind t unnel verificat ion of a top mount ed S shaped inlet ut ilizing t he diverterless
TAN Hui jun Born in 1975, he received Ph. D. degree fro m Nanjing U niversity of Aeronaut ical and Astronautics in 2003, and then became a teacher there. His research inter ests focus on CF D and exper imental investigations of internal flo w. T el: ( 025) 84892203 2415, E mail: tanzi109@ 263. net
No. 2
CH INES E JO U RN A L O F AER ON A U TICS
M ay 2004
Design and Wind Tunnel Study of a Top mounted Diverterless Inlet T AN Hui jun, GU O Rong wei ( I nternal Fl ow Research Cent er , Nanj ing Uni versity of A eronautics and A str onautics , Nanj i ng Abstract:
210016 , China)
Combined w ith a U A V o f the shape like Global Hawk, a new inlet is advanced to obtain
hig h performance in both Radar Cross Section( R CS) and aerodynamic drag. Efforts are made to achieve t his goal such as ado pting a top mounted inlet configuration, ut ilizing the diverterless technique and putting forw ar d a new shape of entrance. A design method is brought for ward and v er ified by wind tun nel tests. Results indicate: ( 1) Despite the negativ e effect of the fr ont fuselage and the absence of the conventional boundar y diverter , t he performance of the top mounted diverterless inlet advanced here ( M a: 0 50 0 70,
:- 4 6,
> 0 975) is equivalent to that o f convent ional S shaped inlet with di
verter; ( 2) T he integration of the inlet w ith the fuselage is realized by t he utilization of a special inlet section and t he div erterless technique, w hich disposes t he whole inlet in the shield of the head of UA V, improving the drag char acteristics and the stealthy performance of the aircraft; ( 3) T he bump which is equal to the local boundary layer thickness in height can diver t the boundary layer effect ively. As a re sult, no obv ious low total pressure zone is found at the out let of the inlet; ( 4) Accor ding to the ex peri mental results, negative angle of attack is favorable to the total pressure recov er y and positive ang le of at tack is favorable to the total pressur e distortion, while y aw brings bad effects on bo th; ( 5) T he design of cowl lip is of great impo rtance to the inlet performance at y aw , therefore, further impro vement of the inlet performance will rely on the lip shapes of the cowl chosen. Key words:
top mounted inlet; div erterless inlet; unmanned air vehicle; design; w ind tunnel test
一种无人 机用背负式无附面层隔道进气道的设计及实验验证. 谭慧俊, 郭荣伟. 中国航空 学报( 英 文版) , 2004, 17( 2) : 72- 78. 摘 要: 结合一类似于 全球 鹰! 的无 人侦察 机外形, 对一 种新型 高隐身 低外阻 进气道进 行了如 下 设计: 采用背负式布局方案, 使用无隔道技术, 并提出了一 种新的进口 截面形状。加 工了风洞 试验 模型并开展了验证性风洞实验研究工作。结果表明: ( 1) 尽管受到 机头遮蔽 的不利影响, 且没 有采 用传统的附面层隔道, 所给出的背负式无隔道 进气道方 案性能( Ma: 0 50~ 0 70,
:- 4 ~ 6,
>
0 975) 与常规的有隔道 S! 弯进气道相当; ( 2) 特殊进口截面形 状及无附 面层隔道技 术的采用 将进 气道与机身有机地融为一体, 使进气 道整体 都处于 飞行器头 部的遮 蔽之中, 这 有利于改 善飞行 器 的阻力特 性和隐身性能; ( 3) 进气道出口截面上未发现因附面层吸 入而造成 的低总压区, 这说 明高 度与当地附面层厚度相当的进口鼓包能有效地隔除附面层中能量较低的气流; ( 4) 研究范围内, 负 攻角对进 气道的总压恢复系数有利, 正攻角 对周向 畸变指数 有利, 而侧 滑角则 对两者均 有着不 利 影响; ( 5) 唇口的设 计对进气道的侧滑角性能有着重要影响, 进气道性能 的进一步提高应考虑唇口 设计的改进。 关键词: 背负式进气道; 无隔道进气道; 进气道; 无人机; 设计; 实验 文章编号: 1000 9361( 2004) 02 0072 07
中图分类号: V211. 3
文献标识码: A
T op mounted inlets have advant ages over con vent ional inlet conf igurat ions as leav ing more space
creasing st ealt hy capabilities t o up looking radar, and prevent ing foreign objects from ing est ing int o
underneat h t o place missiles or ot her weapons, pro viding bett er environments for g round crew , in
eng ines[ 1] . T hey are quit e common on commercial aircraf ts w hen t hree engines are desired like on t he
R eceived dat e: 2003 05 19; R evision received dat e: 2004 02 10
M ay 2004
D esign and W ind Tunnel St udy of a T op M ount ed Divert erless Inlet
# 73 #
M cDonald Doug las DC 10, MD 11. However, t he complex aerodynamics issues, especially at high an g les of att ack, impeding t he top mounted inlets∀ w ide use on milit ary aircraft s, and at present t hey are mainly applied t o UCAVs and bombers such as
Fig. 1
Sketch of the fuselag e
Global Hawk, X 45, X 47 and B 2. In t he fift ies of last century , the explorat ions of the divert erless t echnique w ere init iat ed by re
2
Design of t he Inlet
searchers in NACA Lew is Laborat ory [ 2] . But t he
T he bulge at the front fuselage shadows t he
diverterless technique did not get furt her at tent ions
top mounted inlet and is believed to reduce t he
unt il its recent successful applicat ion in X 35. T his unique technique, eliminat ing t he divert er and ot h
front al aspect Radar Cross Sect ion ( RCS ) of t he
er f eat ures associat ed w ith boundary lay er m anage
inlet . H ow ever, it is a great challenge t o design t he inlet because t he bulge speeds up the development
ment, is realized by desig ning a three dimensional
of t he boundary layer, resulting in the apert ure of
surface, or bump, f ront of the inlet apert ure, w hich can decelerat e part of t he com ing f low and
the inlet mostly in t he boundary layer. In order t o
create a pressure gradient that pushes the boundary
solve the problem mentioned, t he Global Haw k adopt s hig her boundary diverter, w hich lift s t he
layer aw ay from the inlet . T hus, t his technique is
inlet to ameliorate the f low condit ion in front of t he
able to reduce t he overall w eight of t he aircraf t, cut dow n t he front al area of the aircraf t[ 3] , improve
aperture. But t his approach w ill increase RCS and the aerodynamic drag of t he aircraf t. Hence, a top
the drag characterist ics and enhance t he st ealt hy
mount ed inlet int egrated w it h the f uselage is
capabilit ies by decreasing corner reflections of radar
brought forward to dispose t he w hole inlet in t he
w aves. According to dom est ic open literat ure, lit t le
shield of t he bulge. T he diverterless t echnique is u t ilized in this design t o deal w it h the t hick bound
w ork has been made on t he top mount ed inlet or
ary flow and t o obt ain bet ter conformance. Addi
the divert erless technique, let alone t he combina
t ional endeavours are also m ade to acquire f avorable
t ion of both. Integ rated w ith a UAV of t he conf ig urat ion like Global Hawk, a top mount ed S shaped
inlet perf orm ance. T hough the conventional design method[ 4] of
inlet utilizing the diverterless t echnique is dev ised
S shaped inlet is still used here, a creat ive cross
and verif ied by w ind tunnel t ests in t his study.
sect ion is come up w ith, w hich conveniently inte
Description of the Forebody
g rate t he inlet w it h t he f uselage and the bump em ployed by t he divert erless t echnique. T he proce
F ig . 1 demonst rat es t he f orepart of an aircraf t,
dure of t he inlet design is described in det ail in t he
1
w hich resem bles t he Global Haw k, the UAV of
follow ing.
U S Air Force. T he dist ance from t he apex of t he fuselage to t he leading edg e of t he inlet is L . As
2. 1
shown in t he f ig ure, a bulge which is 0 06L in
as lengt h, offset and exit area are subject t o geo
height rests on the top side of the fuselage. Because
metric const raints of t he fuselage and t he engine,
the w ing s are below and far aw ay f rom the inlet , it is obvious t hat t hey have litt le inf luence on the per
they are alw ays know n w hen one set out to design a inlet. According to a selected area rat io, t he en
f orm ance of t he inlet at low incidence. Accordingly
t rance area of t he inlet can t hen be decided. T he
the design and w ind t unnel test s of the inlet in t his paper have not taken the eff ect of the wings int o
paramet ers used in t his paper are summarized in T able 1, t wo of w hich are normalized by t he diam
account .
eter of the ex it .
General geometrical parameters of the inlet F or some general parameters of the inlet such
# 74 #
T A N Hui jun, GU O Rong w ei
Table 1 General geometrical parameters of the inlet
2. 4
2. 2
1 6
1 3
1 5
Design of the entrance shape T he reason w hy t he convent ional cross sec
t ions like ellipse or racew ay shape can not be used
Design of the diffuser T he cross sectional area ( A i ) at arbit rary sta
Lengt h/ Diameter O f fset / Diameter A rea rat io Equivalent angle/ ( ) 5 0
t ion of the diff user can be obt ained by its entrance area, exit area and area dist ribution. Ref erring t o Fig. 2, it is not diff icult t o deduce t he mat hemat i cal expression of any cross sectional area
here is t hat the diverterless technique requires t he
Ai =
integrat ion of the entrance section w it h t he bump, w hile neither of t hese can sat isfy this requirement . T he new ly dev ised cross sect ion is pict ured in F ig. 2, w hich consist s of a cirque ( centre ang le , in t ernal radius r and ext ernal radius r + a + b) and four quarter ellipses ( major axis a or b and minor
i
2
( 1)
ai bi ri ,k = ,k = and assum ing ci 2, i c i 3, i ci these rat ios and i are know n, Eq. ( 1) can be rew rit ten as Lett ing k 1, i =
ci =
its, one is t hat it can be int egrated w ith the bump and t ransformed to the shape of t he engine f ace
+
easily, and t he ot her is that it unit es t he inlet w ith aerodynam ic drag and low observ abilit y.
!a ici !bic i i + + !( r i + a i + b i ) 2 2 2 360 - 360!r i
ax is c) . For t his configuration t here are t wo mer
the fuselage t o the ut most deg ree achiev ing low
CJA
i
360
Ai M
M= !
k l , i + k 2, 2
( k l , i + k 2, i + k 3, i ) 2 -
i
360
i
k 23, i
which means t hat the configuration of this sect ion is ex act ly determined. k 1, i , k 2, i , k 3, i and i vary smoot hly and monotonously from k 1, in , k 2, in , k 3, in and
in
to k 1, out, k 2, out, k 3, out and
out
when t he sta
t ion moves from the ent rance to the ex it . T he val ues of k 1, in , k 2, in , k 3, in and in det erm ine the shape of t he entrance section, which are select ed 1 0, 2 5, 2 5 and 110 F ig. 2 Sketch of the entrance section
2 3
Centerl ine and area distributions T he centerline and t he area distribut ions are of
g reat significance to the performance of the inlet. Centerline shapes determ ine t he turning of the in t ernal f low and t hereby decide t he transverse pres
individually in this st udy.
While determining k 1, out, k 2, out , k 3, out and out one should be more careful to guarant tee t hat the exit of t he dif fuser is an ex act circle. T he values of 1 0, 1 0, 4 0 and 0 are capable to meet the re quirement well and are used here. Result show s that t he boundary condit ions of k 1, i , k 2, i , k 3, i and i
work w ell ( F ig . 3) .
sure gradient and t he secondary flow in t he duct. Area dist ributions det erm ine the diff usivit y and hence control t he streamw ise pressure gradient im posed on the flow. Inappropriate cent erline shape or area dist ribut ion may both lead t o f low separa t ion. Ref. [ 4] provides three poly nomial funct ions w hich are suit able for both centerline shape and area distribut ion. Based upon previous investiga t ions, the cent erline distribut ion w it h a modest
F ig. 3 3 D view of the diffuser
2. 5
Design of the bump
turning and t he area dist ribution w it h a rapid
T he success of t he diverterless technique w ill be largely dependent on t he design of t he bum p. A
change near t he ex it are used in this study.
well designed bump should ex ert its f unct ions of
M ay 2004
D esign and W ind Tunnel St udy of a T op M ount ed Divert erless Inlet
# 75 #
both guiding and pushing to keep the boundary lay
the innovat ive cross sect ion of the inlet , an int egra
er aw ay f rom t he inlet by elaborat e surf ace design.
t ion of the inlet and t he fuselage is obt ained w hich
T he effect of guiding depends directly on t he bump it self and operat es on t he low er part of t he bound
minimizes t he inlet∀ s heig ht and t hereby brings t he whole inlet into the shadow of t he bulg e of t he
ary layer flow . But the ef fect of pushing appears
fuselage.
not so direct and is det ermined by the pressure gra dient formed by t he bump act ing on t he upper lay er . Wit h a crescent shaped cross sect ion and a polynomial vert ical section line, t he bump is ap prox imate t o a cone shaped surf ace, w hich is t an gent to t he fuselage and t he inlet at it s boundaries. Computational result s indicate t hat bet ter perf or mance of the inlet can be achieved by a bump of F ig. 5 Enlarg ed view of the inlet∀ s forepart
w hich t he max imum heig ht is equivalent to t he thickness of the local boundary layer ( evaluat ed by
In addit ion, t he junct ion of t he inlet and t he
turbulent boundary layer of flat plate ) and t he
eng ine ( w it h center body) is designed using con
lengt h is about five times the max imum height . 2. 6 Design of cowl lips
st ant area dist ribut ion. At the same t ime, t he de sign of outside surface of the inlet is also conduct
In general, lips of subsonic inlets are of impor t ance t o the performance of t he inlet. And, w it hal,
ed.
3
in t his paper it act s as an addit ional measure assist ing the bump t o push boundary f low aw ay from t he inlet. A lip scheme wit h variable section line and [ 5]
v ariable lengt h is adopt ed. NACA 1 85 100 lip
is
Experimental Verification
In order to obtain t he aerodynamic perf or mance of t he new ly designed inlet and verify t he design approach, an experiment al inlet model is de
at top point of t he ent rance ( T point in F ig . 2) forming a local ∃dif fuser∀ ( F ig. 4) to acquire high
signed and manufactured and t hen test ed in a wind tunnel.
er pressure in reg ion A t han t hat of t he outside. Sym met ry airfoil is at t he intersect ion point w ith
3. 1
the bump ( S point in F ig . 2) to prevent perf or mance deterioration at yaw . T he ut ilization of vari able lengt h lip gets a forw ard sw ept cow l which co
Inlet model and test tunnel
T he complet e t est model consist s of fuselage, top mounted inlet, eng ine cent erbody, pressure rakes and conic flow plug as illustrated in Fig. 6.
operat es w ith the local ∃ diffuser∀ and ensures t he
Due to the constraints of w ind t unnel blockage and the m easurement of internal f low , t he w ings are
majorit y of the boundary layer f low spilling out at the aft notch.
not taken into considerat ion. T he t est model has a tot al length of 1000mm and the diameter at t he outlet of the inlet, where the measurement of t he pressures is made, is 40m m.
Fig . 4 F eature o f the upper lip and the inlet bump
F ig. 6 Sketch of the test model
F ig. 5 gives an enlarg ed view of t he inlet ∀ s forepart. In virtue of t he divert erless technique and
An eight leg rake is inst alled at t he out let of the inlet for measurement of st eady tot al pressure.
# 76 #
T A N Hui jun, GU O Rong w ei
CJA
Each leg has five probes, providing a tot al of 40 lo
wards. T he possible reason may be that at low
cat ions. Four st at ic pressure taps are locat ed on t he
mass f low rat io more low energy flow near t he
inner surface of the inlet equally along the circum
fuselage enters t he inlet relat ive t o t he total mass
f erence and in a plane cont aining the face of the t o t al pressure probes. T he m ass f low plug , placed at
flow captured. ( 2) Ef fects of free st ream Mach number
the model ex it and driven by an electric mot or,
F ig. 8 shows the variat ion of tot al pressure re
prov ides flow cont rol for t he measurement s at dif f erent mass flow rat ios.
covery coeff icient w it h f ree st ream M ach number. It can be noted t hat the recovery coef ficient is in
T he experiment al study is accomplished at
sensitive t o M ach number up to 0 7 and remains
NH 1 high speed wind t unnel of NUAA. T he
above 0 975 st ably. T his indicat es that , despite
Reynolds number of t he w ind t unnel ranges f rom 1 2 % 107 to 1 77 % 10 7 . During this t est, t he free
the neg at ive impact of t he bulge and t he absence of the boundary diverter, the recovery coeff icient of
st ream Mach number varies f rom 0 5 t o 0 8, ang le
the inlet tested is equivalent t o that of convent ional
of at tack from - 4 t o 6 and yaw from 0 to 2 . T he maximum blockage of the t est sect ion is 4 3%
S shaped inlet w it h diverter. How ever, w hile t he Mach number is greater than 0 7, t he recovery co
w hile t he inlet model is inst alled.
eff icient t akes on a sharp decline. T his t rend is
3 2
probably aroused by a local shock occurring at t he
Experimental resul ts
T he performance of the inlet is evaluat ed by the t otal pressure recovery coef ficient ( ) and cir
back of t he bulge, w hich thickens rapidly t he boundary layer and det eriorat es t he flow condit ion
cular t otal pressure distort ion ( ∀ ) . In the f ollow
front of the inlet. It also means that t he fuselage
ing text , t he eff ects of mass flow ratio, f ree st ream M ach number, angle of at tack and yaw are illus
conf igurat ion like Global H aw k is not suitable for high subsonic flight .
t rat ed. Cont ours of tot al pressure at the outlet of the inlet are also presented. ( 1) T ot al pressure recovery coef ficient at dif f erent mass flow rat ios In Fig. 7, t he tot al pressure recovery coeff i cient is plott ed as a funct ion of mass f low rat io at 0 ang le of att ack and yaw . It can be seen t hat , at first , w it h t he descent of mass f low rat io t he recov ery coeff icient ascends due t o the reduction of t he frict ion loss. But when the m ass flow rat io goes dow n further the recovery coeff icient t ends dow n
Fig . 7 T o tal pressure r ecovery coefficient v ersus mass flo w ratio( = 0 , #= 0 )
Fig . 8 T o tal pressure recovery coefficient versus free stream M ach number
Fig. 9
Cir cular distortion coefficient versus free stream Mach number
M ay 2004
D esign and W ind Tunnel St udy of a T op M ount ed Divert erless Inlet
# 77 #
For all test M ach numbers, t he circular total
trance of the inlet is lower t han that of free st ream
pressure distortion keeps below 1. 2% , as show n in
because of the influence of the upper fuselage,
F ig. 9. It means t hat the dist ribut ion of tot al pres sure losses is nearly constant along circular direc
which results in a f ew effect on circular distort ion. For t he case of posit ive angle, because the count er
t ion.
rot at ing vort ex pair developed by t he up w ash f low around t he fore fuselage sweeps off part of t he low
( 3) Eff ect s of incidence T he model is tested at six ang les of at tack, of w hich t he range is f rom - 4 t o 6 . Fig. 10 and F ig. 11 present t he ef fects of ang le of at tack on t o t al pressure recovery coef icient and circular total pressure dist ortion respectively. It can be observed from t hese t wo f igures that negative angle of at tack benefit s recovery coef ficient and aff ect s distort ion slightly, w hile posit ive angle of at tack affects re covery coefficient slight ly and improves distort ion obv iously. For t he case of negative ang le, it can be
energy flow in front of t he inlet , t he curve of t he tot al pressure recovery goes up a bit f rom 0 t o higher incidence and hence a remarkable drop of circular dist ortion occurs. ( 4) Ef fects of yaw T he ef fects of yaw on t otal pressure coeff icient and circular tot al pressure distort ion are illustrated in F ig. 12 and Fig. 13. It can be seen that 2 y aw has lit t le ef fect on recovery coeff icient but increases
explained as follows: on t he one hand, w it h nega
dist ortion distinctively ( still less t han 2% ) . More over, w ith t he increase of t he Mach number, t he
t ive angle of att ack rising, the ent rance of t he inlet is exposed t o f ree stream more and more, result ing
performance of the inlet is det eriorating , in part ic ular, at hig h M ach number, which can be associat
in more main flow captured and thereby t he in crease of tot al pressure recovery coef ficient; on t he
ed w it h unfavorable f low occurring near the lip at
ot her hand, the local flow ang le of at t ack at the en
higher M ach number. Consequently, t he improve ment of perf orm ance at y aw w ill lies on the melio ration of cow l lips.
F ig. 10
T otal pressure recovery coefficient versus angle of attack
Fig . 12 Effects of yaw on to tal pressure recovery co efficient
Fig . 11 Circular total pressure distortion co efficient versus ang le of attack
F ig. 13 Effects of yaw on circular distortion coefficient
# 78 #
T A N Hui jun, GU O Rong w ei
CJA
( 5) T otal pressure contours
technique are reported. During the design process,
F ig . 14 demonst rat es the t otal pressure dist ri
the int eg rat ion of the inlet w it h t he fuselage is real
butions at t he outlet of t he inlet w hile Ma = 0 7,
ized by t he utilizat ion of a special inlet sect ion and
= 0 , #= 0 . T he t ot al pressure cont ours make a feature of high m agnit ude and uniform distribut ion
elim inat ion of convent ional diverter, w hich dispos es t he whole inlet in the shadow of the bulge im
and no obvious low tot al pressure reg ion ex ists. All
proving t he drag characterist ics and t he st ealt hy
of t hese suggest t hat t he diverterless t echnique used in this paper can ef fectively push the low energ y
performance of t he aircraft . Wind tunnel tests indi cat e t hat the perf orm ance of t he inlet advanced here is equivalent to t hat of convent ional S shaped inlet
flow aw ay from t he inlet , w hich is validated by CFD result s ( F ig . 15) [ 6] .
wit h diverter, so t he diverterless t echnique adopted is successful.
References [ 1]
V ooren A , Kewenter H, Edefur H . T op mount ed air inlet for supersonic & transonic milit ary aircraft [ OL] . ( 2001 05 16 ) ht t p: / / w w w . ave. kt h. se/ education/ msc/ cours es/ 4E1232/ 2001/ aero2/ T M I R eport . pdf.
[ 2]
S imon P C, Brow n D W, Huf f R G . Performance of ext ernal compression bump inlet at M ach numbers of 1 5 t o 2 0 [ R ] . N A CA RM E56L19, 1957.
[ 3]
M cFarlan J D. Joint strike f ighter diverterless supersonic inlet [ O L ] , ( 2000 11 15) . ht tp: / / w w w. codeon emagzine. com/
F ig. 14 Contours of total pr essure recov er y coefficient
archieves/ 2000/ art icles/ july 00/ divertless
at the outlet [ 4]
1. ht ml.
Lee C C, Boekicker C. Subsonic dif fuser design and perf or man ce f or advan ced fight er aircraft [ R ] , A IA A 85 3073, 1985.
[ 5]
R ichard J R . A n invest igation of several NA CA series inlets at M ach number from 0 4 to 1 29 for mass f low rat io near 1 0 [ R ] . N AS A TM X 3324, 1975.
[ 6]
谭慧俊, 郭荣伟. 背负式无 隔道进气 道的方案 设计与性 能 研究[ R ] . 南京航空航天大学科技报告, R 0202018, 2002. T an H J, Guo R W. Design an d performance st udy of a top mounted diverterl ess inlet [ R ] . T echnical Report of N U AA , R 0202018, 2002. ( in Ch inese)
Biography: Fig . 15
Streamlines released from points 2mm away from fuselage
4
Conclusions
Design and w ind t unnel verificat ion of a top mount ed S shaped inlet ut ilizing t he diverterless
TAN Hui jun Born in 1975, he received Ph. D. degree fro m Nanjing U niversity of Aeronaut ical and Astronautics in 2003, and then became a teacher there. His research inter ests focus on CF D and exper imental investigations of internal flo w. T el: ( 025) 84892203 2415, E mail: tanzi109@ 263. net