NASA Technical Memorandum 103170

_ Experimental and Analytical Study of

Close-Coupled Ventral Nozzles

for ASTOVL Aircraft

Jack G. McArdle and C. Frederic Smi.th

Cleveland, Ohio

Prepared for the

International Powered Lift Conferenc_¢ ._

sponsored by the Royal Aeronautical Society

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(NA_-A - ] M-103i70) EXp_:_IMFNTAL ANO

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NOZ_LFS FOP, A_TF!,VL AIPC_AFT (NASA) 16 p

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N_I07 0290723

EXPERIMENTAL AND ANALYTICAL STUDY OF CLOSE-COUPLED

VENTRAL NOZZLES FOR ASTOVL AIRCRAFT

Jack G. McArdle

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio, U.S.A

and

C. Frederic Smith

Sverdrup Technology, Inc.

Cleveland, Ohio, U.S.A

9"

Abstract

Flow in a generic ventral nozzle system

was studied experimentally and analytically with a block version of the

PARC3D computational fluid dynamics

program (a full Navier-Stokes equation

solver) in order to evaluate the program's ability to predict system performance and internal flow patterns.

For the experimental work a one-thirdsize model tailpipe with a single large

rectangular ventral nozzle mounted

normal to the tailpipe axis was tested

with unheated air at steady-state pressure ratios up to 4.0. The end of the

tailpipe was closed to simulate a

blocked exhaust nozzle. Measurements

showed about 5%-percent flow-turning

loss, reasonable nozzle performance

coefficients, and a significant aftward

axial component of thrust due to flow

turning more than 90 °. Flow behavior

into and through the ventral duct is

discussed and illustrated with paint

streak flow visualization photographs.

For the analytical work the same ventral system configuration was modeled

with two computational grids to evaluate the effect of grid density. Both

grids gave good results. The finergrid solution produced more-detailed

flow patterns and predicted performance

parameters, such as thrust and discharge coefficient, within 1 percent of

the measured values. PARC3D flow visualization images are shown for comparison with the paint streak photographs.

Modeling and computational issues

encountered in the analytical work are

discussed.

Introduction

Improved short takeoff and vertical

landing (STOVL) aircraft are planned

for possible future development. For

these aircraft the same propulsion system will provide power for lift and

hover as well as for supersonic horizontal flight. The propulsion system

must have high-thrust-to-weight-ratio

engines; efficient gas ducting and

thrusters; large, reliable valves and

seals; and integrated engine and flight

controls. To develop the required

technology for these aircraft, the NASA

Lewis Research Center has established

active programs in mission analysis,

integrated controls, lift thrusters,

and hot gas reingestion in hover

flight.

When the STOVL propulsion system is

operating in the lift mode, the rear

jet nozzle will be blocked, and valves

will be opened to duct engine exhaust

gases to two or more thrusters directed

downward. In many proposed configurations one of the lift thrusters will be

a ventral nozzle. A typical arrangement is sketched in Fig I. The ventral

nozzle draws mixed core and fan gases

from the engine tailpipe through a

valve and opening having no inlet turning vanes. The ventral nozzle size and

location are chosen to balance the

pitching moment from thrusters forward

of the aircraft center of gravity during hover flight. Close coupling is

necessary because the valve and the

nozzle must be wholly mounted within

the fuselage. The ventral nozzle also
EJECTOR

EXHAUST\ FI]OENGINE J / NOZZLE.3

/

L-VENTRAL

VALVEAND

NOZZLE

FIG I. - PROPOSEDSTOVLPROPULSIONSYSTEMWITHVENTRALNOZZtE,

ON E7DAIRCRAFTMODEL.

may be vectored to provide trim or

pitch control.

Early work on STOVL deflectors was

reported in Refs I and 2 and included

experiments examining the internal flow

field and performance of model turbofan

engine and ventral nozzle configurasteady-state ratios of tailpipe to

ambient pressure up to 4.0, internal

paint streak flow visualization photographs, and CFD flow visualization

images for direct comparison with the

experimental photographs at a pressure

ratio of 3.0.

Apparatus and Instrumentation

Test Stand

The model was tested on the Powered

Lift Facility (PLF) at the NASA Lewis

Research Center. The PLF is a unique

outdoor stand designed to measure

simultaneous axial (thrust), normal

(lift), and side forces up to 25 000 Ib

(ii0 kN). The stand was supplied with

unheated air from the central system at

150-psi (1000-kPa) pressure. Airflow

tions.

tional fluid dynamics (CFD) analyses of

ventral flow were found in the literature. In an ongoing Lewis program

ventral nozzle performance is being

studied experimentally and analytically. A generic model tailpipe (about

one-third of full size) having a single

large, rectangular ventral nozzle

mounted normal to its axis was built

and tested with unheated air. A blind

flange at the end of the tailpipe

simulated a blocked exhaust nozzle.

Analytical performance of the same

tailpipe and ventral nozzle configuration was modeled and studied with a

CFD code called PARC3D, using Cray

computers. PARC3D solves the threedimensional, compressible-flow NavierSto_es equations and includes an

algebraic turbulence model. A block

version of PARC3D was used to simplify

construction of the chosen computational grid and to reduce the computer

requirements for a fairly complex geometry. Two grids for the same configuration were generated to evaluate the

effect of grid density on the numerical

results.

However, no reports of computa- was controlled by a valve in the facility inlet line. Flow rate was measured

with an ASME long-radius nozzle. The

model mounted on the PLF stand is shown

in Fig 2. The model was mounted with

the ventral nozzle discharging upward

for operational convenience.

The major objectives of the work

reported in this paper were to expand

understanding of ventral flow turning

by appropriate tests, and to evaluate

the ability of the chosen computational

grids and CFD code to predict the

experimental performance. The results

are shown in performance plots for

Model

A sketch of the model tailpipe and ventral nozzle is shown in Fig 3. An

uncomplicated design was chosen in

order to simplify the computational

grid generation. The tailpipe was

13.5 in. (34.3 cm) in diameter, which

is about one-third the size of many

current military engines. The model

was connected to the 24-in. (61-cm)

diameter facility mounting flange

through a conical-plus-cylindrical

transition section. This section conFIG2. - MODELMOUNTEDON POWEREDLIFTFACILITYlESTSTAND.

ORIGINAL PAGE

BLACK AND '._HJTEPHOIOGP, AP_H
_G.q-[N. X 11.7-IN. (16.2-CM x 29,7-CM)

\ VENTRAL NOZZLE

L OPTIONAL FLOW

VISUALIZATION

PLATE

SECTION A-A

---27 IN.(68,6 ' VISUALIZATION

PLATE

A

_MODEL TAILPIPE LENGTII, q8,7 IN.(12q C_)--

ORIGINAL PAGE IS

OF' POOR QUALITY

A

r BOUNDARY

I/ LAYER TRIP T- FLOW

N

•,'-,,--STATION GB i FSCREEN AND FLOW _STRAIGHTE ER

STATION G (21J _'_ -STATION S ;/ t STRAIGHTENER _, J

 -BLINOTOTAL P ESSORES -- II I OTOTAL/ ' t

y FLANGE II |PRESSURES) /' I // \_ ,

.... !%,....

- OPTIONAL FLOW __Z=:

--TRANSITION SECT[ON_

FACILITY

MOUNTING

FLANGE

FIG 3, - EXPERIMENTAL MODEL OF TAILPIPE/VENTRAL NOZZLE.

tained two honeycomb flow straighteners

and one fine-mesh screen, just ahead of

the model, designed to provide uniform

inflow. The screen was 14 mesh by

0.009-in. (0.23-mm) wire, and on the

basis of results reported in Ref 3, was

relied on to reduce free-stream turbulence intensity to less than 1/2 percent. In addition, a toothed metal

strip protruded 0.06 in. (1.5 mm) from

the wail to ensure that the inflow

boundary layer was turbulent. A blind

flange was located two tailpipe diameters downstream of the ventral opening

to simulate a blocked exhaust nozzle.

The rectangular convergent ventral nozzle was mounted on a duct 0.89 tailpipe

diameter long (measured from the tailpipe centerline). The intersecting

edges of the ventral duct and tailpipe,

at the ventral cutout, were not

rounded. These "square" edges were

intended to represent a worst-case

design condition.

An optional flow visualization plate,

0.06 in. (1.5 mm) thick, could be

mounted on the vertical centerline

(a plane of symmetry) of the model over

the ventral opening. The plate

extended one tailpipe diameter downstream of the opening to capture flow

patterns in the tailpipe. The plate

Instrumentation

The model instrumentation stations are

shown in Fig 3. Tailpipe pressure (P5)

was measured with five total-pressure

tubes on each of four equally spaced

rakes at station 5. Ventral nozzle

inlet pressure (P6) was measured with

24 total-pressure tubes (20 tubes

arranged uniformly in the duct plus 4

corner tubes) at station 6. The wall

pressure at several locations on the

model was measured with static-pressure

taps. Ventral nozzle exit flow conditions were measured by probes driven

across the nozzle just downstream of

the exit plane by an electric actuator.

For one test a three-tipped totalpressure probe was used to map the

pitot-pressure distribution. For

another test a calibrated conical probe

was used to measure total pressure, as

well as flow angles relative to the

probe tip, as functions of the measured

tip and surface pressures. Airflow,

forces, air temperature, and ambient

conditions were measured by facility

instrumentation systems using calibrated load cells and other conventional transducers.

Procedure

was not installed for performance tests Performance Tests

or flow visualization photographs of

the tailpipe and ventral duct walls. After force system tares had been
obtained, steady-state thrust and airflow performancewasmeasuredat several ratios of tailpipe to ambient

pressure (PRS)up to 4.0. The data

were recorded on the central laboratory

system and batch processed on a mainframe computer.

Exit Surveys

The flow-angle probe and total-pressure

rake data were obtained at selected

locations in the exit flow (station 6B)

at PR5 _ 3. The data were processed

along with the performance test data.

Flow Visualization

After the optional flow visualization

plate had been installed, dabs of

thick, oily paint were applied to the

plate with a syringe in a grid-like

pattern. In order to minimize transient flow effects, airflow was started

quickly, held at PR5 - 3 for 1 min, and

then quickly shut down. The flow

caused the paint to run along streamlines, and the resulting streaks provided a clear picture of the flow

pattern. A similar procedure was followed without the plate installed to

obtain flow visualization photographs

at the ventral duct walls.

Computational Fluid Dynamics

A major objective in these studies was

to evaluate the ability of a chosen

computational grid and CFD code to predict the internal flow patterns and

overall performance of a ventral nozzle

system. To this end, the PARC3D code

was chosen because it is familiar to

NASA Lewis researchers and because it

is known tO be applicable to flow problems of the same general type. The

same tailplpe and ventral nozzle configuration used for the experimental

work was modeled as described in the

following paragraphs.

PARC3D Code

The PARC3D code was originally developed at the NASA Ames Research Center

(Ref 4) to analyze external flows.

It solves the three-dimensional,

compressible-flow, Navier-Stokes equations in curvilinear coordinates and

includes the Baldwin-Lomax (algebraic)

turbulence model (Ref 5). The code

later was modified at the U.S. Air

Force Arnold Engineering Development

Center (Ref 6) for use with internal

flows. The equations of motion are

solved by using Beam and Warming's

approximate factorization scheme.

Pulliam's scalar pentadiagonal transformation is applied to uncouple the

equations of motion.

A block version of the PARC3D code was

used in this study. The block version

allows the computational grid to be

broken into two or more blocks to

simplify modeling. Trilinear interpolation is used to exchange computational information between modeling

blocks (Ref 7). As an additional

advantage, most of the information

required in the core memory is associated with the current block, thus

reducing computer storage needs.

Computational Grid

Because the experimental configuration

had a plane of symmetry, only one of

the symmetric halves was modeled. The

grid model was broken into two blocks.

The tailpipe made up one block, which

was modeled by using an O-grid. The

O-grid consisted of concentric circles

parallel to the tailpipe surface and

radial lines normal to the surface.

The ventral duct and nozzle made up the

other block and were modeled by using

an H-grid. The H-grid consisted of

lines parallel and perpendicular to the

walls. This approach provided bodyconforming grids for each block. A

"wire" diagram of the grid model is

shown in Fig 4 along with a photograph

of the experimental model for

comparison.

Both grids were generated algebraically

by the INGRID3D code (Ref 8). The

grids were stretched by hyperbolic

functions in order to pack grid points

near the walls and the center of the

0-grid. Two grids of the same model

were made to evaluate the effects of

grid density on the predicted flow

field and numerical results. The first

grid, called herein the "coarse" grid

and shown in Fig 5, contained 31 875

points in each block (51x25x25 points
FLOW

r VENTRAL DUCT

L

r PLANE OF I _'___'I_

" _ 41 FLOW

'_SYI_ETRY

i__ "_ TAILPIPE " "

/ BL_KED

END

(a) ANALYTICAL "WIRE" DIAGRAM. ONLY ONE-HALF IS USED FOR CFD:

FOR CLARITY MANY OF THE GRID LINES ARE NOT SHOWN.

VENTRAL NOZZLE AIRFLOW INLET

(b) EXPERIMENTALMODEL.

FIG _, - VENTRALSYSTEM CO_IGURATION.

ORIGINAE PA_E

BLACK AND WHITE PHO_OGRAPH

_VENTRAL DUCT

/ AND NOZZLE

,'

._ _ UPSTREAM

DOWNSTREAM WALL-'"_" WALL

'_ TA[LPIPE

(a) PLANEOF SYMMETRY (xz-PLANE).

_VENTRAL DUCT

AND NOZZLE

PLANEsYMMETRyOF _L"

- SIDE WALL

_VENTRAL DUCT

/ AND NOZZLE

DOWNSTREAMWALL_,_-UPSTREAM WALL

BLOCKED _: :_ I , I ! _: __

"-TAILPIPE

(a) PLANE OF SYMMETRY (xz-PLANE).

i _ VENTRAL DUCT fli AND NOZZLE

SIDE WALL

PLANEOF

SYI_'TRY _

_ TAILPIPE

(b) END VIEW (yz-PLANE).

FIG G. - "FINE" C_UTATIONAL GRID.

FLOW

in the streamwise and other coordinate

directions, respectively). This grid

size was limited by the space available

on the Cray X-MP computer at Lewis but

was suitable to demonstrate the applicability of the PARC3D code to this

problem. So that an adequate number of

points could be provided in the central

region of the ventral nozzle, packing

near the nozzle walls was limited and

the closest grid point was 0.i in.

(2.5 mm) from the surface. Good analytical results were obtained with this

grid.

The second grid, called herein the

"fine" grid and shown in Fig 6, contained 262 701 (101X51XSl) points in

each block. In the ventral duct and

nozzle block the grid was packed so

that the closest point was 0.01 in.

(0.25 mm) from the nozzle surface. In

the tailpipe the closest point was

0.01 in. (0.25 mm) from the wall surface, the same as in the coarse grid.

Comparison of Figs 5 and 6 illustrates

the denser grid packing near the model

walls and near the tailpipe centerline

obtained with the fine grid. Packing

near the centerline minimized the

effects of the "pole" boundary condition, which is discussed in the next

section.
Boundary Conditions

The boundary conditions for this problem are shown in Fig 7. Analysis of

the jet plume outside the nozzle was

not performed because it would unduly

enlarge the computational grid. However, proper modeling of the flow at

the nozzle exit is very important. The

nozzle has a convergent shape, and the

transonic flow at the exit complicates

the boundary conditions. This problem

was resolved by adding a fictitious

divergent section of suitable size to

the ventral nozzle to provide supersonic flow at the exit of the computational domain. Flow properties could

then be extrapolated downstream of the

area of interest, and the nozzle throat

was modeled as if the external plume

were included. Previous work has shown

that this is a good approach to modeling a three-dimensional convergent

nozzle.

A problem also occurs at the center of

the tailpipe O-grid, where the radial

grid lines become coincident. This

marion derivatives (metrics). This

problem was circumvented by using a

"pole" boundary condition at the center

of the O-grid. In this case the flow

properties were averaged around the

adjacent grid line, and these values

were applied to all points along the

innermost grid line, which had a radius

of about i percent of the tailpipe

radius.

Results and Discussion

This portion of the paper first reports

the results of the performance testing

at steady-state ratios of tailpipe to

ambient pressure (PR5) up to 4.0 and

shows the results of the PARC3D computations at PR5 = 2.96. Next, the

internal flow patterns are described by

using the paint streak photographs and

CFD particle trajectory plots. The

particle trajectory plots were produced

by the PLOT3D code (Ref 9) from the

PARC3D numerical results. Finally,

modeling and computational issues

encountered in the analytical work are

discussed.

would cause singularity problems in

calculating the coordinate transforEXTRAPOLATE

I

_----i_I_FICTITIOUS

DIVERGENT

._.\ SECTION

I" _TRUE NOZZLE

EXIT PLANE

irNO SLIP,

ADIABATIC

. _ I NTERPOLATE

POLE _ L _

r NO SLIP,

ADIABATIC

J

PLANEOF SYMMETRY

FIG 7. - BOUNDARYCONDITIONS.TOTALPRESSURE

AND TOTALTEMPERATUREWERESPECIFIEDFOR UPSTREAMPORTIONOF TAILPIPE;ADIABATICONDITIONS

ANDNO SLIPWERE SPECIFIEDFOR DOWNSTREAMPORTION

OF TAILPIPE.

The pressure ratio PR5 was chosen for

presentation of the results because the

tailpipe pressure is most important to

engine operation and must be held constant during ventral system use to keep

the engine running at the same operating point.

Ventral System Performance

The gas stream lost energy in turning

from the tailpipe into the ventral

duct, resulting in decreased average

total pressure. For STOVL aircraft the

flow-turning loss can limit the maximum

available ventral thrust during hover

flight. The loss measured in the model

tested, shown in Fig 8, was about

5% percent when the tailpipe pressure

ratio was greater than 2. The loss

computed by the PARC3D program at

PR5 = 2.96 was only slightly higher.

This magnitude should be typical of

ventral system designs of this general

type and size. Turning vanes or a

rounded shape at the ventral inlet

could reduce the loss but might be

difficult to package in the small

available fuselage space along with a

shutoff valve and other hardware needed
W_

6

G

,,-.-PARC3DRESULT

/

Q

x>--o'-- r"

I I I I I I

2.0 2._I 2.8 3.2 3.6 ILO

TAILPIPE PRESSURE RATIO, PR5

O.

FIG 8. - TOTAL-PRESSURE FLOW-TURNING LOSS.

at the tailpipe opening. The data in

Fig 8 show that the loss was less at

lower tailpipe pressure ratios. This

trend implies that flow-turning loss is

dependent on tailpipe velocity or Mach

number; therefore, the loss should be

lower in other similar configurations

with smaller ventral nozzles. Both

the measured and analytical system

performance are given in Fig 9 and

Table I[i]. For the ventral system

tested, the discharge and thrust coefficients based on tailpipe conditions,

shown by the open symbols in Figs 9(a)

and (b), were lower than typical for a

simple conical nozzle. However, this

effect is due to the flow-turning loss.

The coefficients rose to more normal

levels when computed in the conventional manner, by using the averaged

nozzle inlet total-pressure ratio (PR6)

as shown by the solid symbols. The

PARC3D results and the experimental

data are in excellent agreement. Both

[i] For these results the parameters

are defined as follows: (I) discharge

coefficient, the measured flow rate

divided by the ideal flow rate at the

same inlet conditions and pressure

ratio; (2) total thrust coefficient,

the measured total thrust divided by

the ideal thrust produced by the measured flow at the same inlet conditions

and pressure ratio; (3) referred flow

rate, the measured flow rate times V_

divided by _, where 8 is the flow

total temperature divided by 518.7 °R

and _ is the flow total pressure

divided by 14.696 psi; (4) referred

thrust, the measured thrust divided by

6.

the measured and CFD-predicted forces

show a negative thrust component,

although the ventral nozzle axis was

normal to the tailpipe centerline.

This negative force is interpreted to

mean that the jet had effectively

turned more than 90 °, as plotted in

Fig 9(c). The axial force measured

7 to i0 percent of the ventral nozzle

total thrust (Fig 9(d)); the PARC3Dcomputed force also was about 7 percent. In a STOVL aircraft the axial

force would tend to accelerate the

craft backward but could be overcome by

vectoring the ventral nozzle or opening

some other thruster to produce a counteracting force.

The measured and predicted vehtral system performances are compared in more

detail in Table I. Both the coarseand fine-grid results are in very good

agreement with the measured values. In

some cases the fine-grid values are

slightly closer to the experimental

data, but the coarse-grid computations

0 EXPERIMENTALDATA

[] PARC3DRESULTS

OPEN SYP_OLSSHOW COEFFICIENTSBASED ON

TAILPIPE (STATIONS) CONDITIONS

SOLID SYMBOLSSHOW COEFFICIENTSBASED ON

I'O-- AVERAGEDNOZZLEINLET (STATION6) CONDITIONS

- = ; -

.8 I t .____1 I I 1

(a) DISCHARGE COEFFICIENT.

(b) THRUST COEFFICIENT.

100

w_ _

(c) JET ANGLE.

°ii b i -lO

_ -1, l=__ l I I I

1.6 2.0 2._a 2.8 3.2 3.6 41.0

TAII.PIPEPRESSURERATIO,PR5

(d) AXIAL FORCE EOI_OClENT.

FIG 9. - VENTRALSYSTERPERFORRANCE.

7
TABLE I. - VENTRAL SYSTEM PERFORMANCE

Ratio of tailpipe to ambient

pressure, PR5

Ratio of nozzle inlet to tailpipe

pressure, PR6

Nozzle discharge coefficient based

on P5

Nozzle thrust coefficient based

on P5

CFD results

Coarse

grid

2.97

0.932

0.88

0.96

Nozzle flow rate referred to PS, 22.67*0.2

ib/sec

Nozzle total thrust referred to 870*9

P5, ib

Nozzle axial thrust referred to -88

P5, ib

Fine

grid

2.96

0.935

0.91

0.97

23.32*0.04

Experimental

2.96

0.945

0.91

0.97

23.04"0.12

905_2 896"10

-72 -73

may be quite acceptable for screening

candidate configurations or otherwise

estimating expected performance.

t FLOW FLOW

Flow Behavior

In this discussion the ventral system

is assumed to be installed in an aircraft as illustrated in Fig i. The

results presented here are for a

steady-state tailpipe pressure ratio,

PR5 = 3. The Mach number at station 5

measured 0.33 on the ventral side of

the tailpipe and 0.27 on the opposite

side.

FLOW

(a) COARSE GRID. (b) FINE GR]D.

Flow in ventral duct - In order to

understand the manner in which flow

turned from the tailpipe into the ventral opening, a test was made with the

optional flow visualization plate

installed as illustrated in Fig 3. The

paint streaks resulting from this test.

along with particle trajectories [ii]

computed with the PARC3D code on the

model plane of symmetry, are shown in

Fig i0. The paint streaks on the plate

(c) EXPERIMENT.

[ii] A particle trajectory is the path

a massless particle would take as it

flowed through the ventral system.

(d) PREDICTION.

FIG 10, - TAILPIPE PLAIC-OF-SW'I_I1tY FLOWVISUALIZATION.

FLOW
show that the flow did not turn sharply

but rather split so that some of the

flow made a smooth turn into the opening and some impacted the tailpipe wall

to form a stagnation point slightly

downstream of the opening. The impacting flow then split, with some flowing

back into the ventral opening and some

flowing farther into the tailpipe. In

the tailpipe it circulated in a counterclockwise direction and returned

upstream along the wall opposite the

ventral opening.

The PARC3D results give the same general flow patterns. (The discontinuities in the particle paths at the

tailpipe centerline, Fig 10(d),

resulted because the PLOT3D program did

not recognize the pole at the center

of the O-grid; however, the PARC3D

solution is continuous across the centerline.) The predictions for the

fine-grid solution reveal several smaller vortices that are not present in

the coarse-grid solution. The predicted center of the larger vortex is

farther from the tailpipe centerline

than shown by the paint streaks. The

reason for this difference may be that

the flow visualization plate is a

no-slip flow surface, whereas the trajectories were computed along an

inviscid plane of symmetry. The recirculating flow was moving along almost

all the tailpipe surface at fairly low

velocity. The wall flow could affect

the cooling-air requirements for ventral systems in STOVL aircraft.

PARC3D velocity vectors (not shown

herein) indicate that in the ventral

duct the flow was separated from the

upstream wall. The separated region

was confirmed by measured wall pressures, which were lowest on the

upstream wall. The flow continued to

turn as it moved through the duct.

Continued turning suggests that the

flow condition at the ventral nozzle

inlet could be influenced by the length

and shape of the duct and, in turn,

could affect the total performance of

the ventral system.

Although the duct flow patterns have

been described in two-dimensional

terms, in fact the flow was highly

three dimensional and many secondary

and vortex-like flows existed, as discussed in the next section.

Flow alonK ventral duct walls - For

another flow visualization test the

optional plate on the tailpipe centerline was removed, and dabs of paint

were applied to the ventral duct and

nozzle walls. The next several figures

show paint streak photographs from this

test and corresponding particle trajectories. Although trajectories were

computed only for one of the symmetric

halves, in the following discussion a

mirror image of each trajectory is

assumed to exist in the other half.

The mirror image is included in the

figures to assist in comparing the

experimental and analytical results.

The paint streaks and particle trajectories obtained on the front (upstream)

ventral wall are shown in Fig ii. Both

show that the separated region consisted of twin counterrotating vortices

FIG 11.- FLOWVISUALIZATION VENTRALDOCTFRONTWALL.
that drew flow from the outer corners

and returned it to the tailpipe near

the centerline. Static pressures measured on the front wall were low, indicating high-velocity flow.

Both the paint streaks and the particle

trajectories on the side walls (Fig 12)

showthat the air flowed out of the

ventral duct and toward the vortices at

the front wall. In the front corners

the air turned along the wall toward

the nozzle.

Twodistinct flow fields can be seen on

the rear (downstream)ventral duct wall

(Fig 13). Air that entered the duct

opening from the downstreampart of the

tailpipe (Fig i0) set up a recirculation region near the tailpipe cutout

revealed by the paint streaks aimed

back to the cutout. This flow mayhave

beenweakbecausethe paint collected

on the sharp edgeof the cutout. The

recirculation was also seen in the

velocity vector field computed by the

PARC3D program (not shown here) and in

FIG I_-' FLOW VISUALIZATIO_ ON VENTRAL DUCT REAR WALL.

FIG 12. - FLOW VISUALIZATION ON VENTRAL DUCT SIDE WALL.

the particle trajectories. The main

flow, however, went directly along the

wall toward the ventral nozzle.

Wall pressure - Experimental and PARC3D

static pressures on the tailpipe wall

are shown in Fig 14. The measured data

clearly show the lower wall pressures

caused by flow entering the ventral

duct. The pressures predicted by the

PARC3D code generally are about 2 percent greater than the measured pressures. The code appears to be less

accurate near the downstream edge of

the tailpipe opening, where the flow

direction suddenly changed at the stagnation region on the tailpipe wall (see

Fig i0).

Nozzle exit flow - A flow-angle survey

was made in the nozzle exit flow with a

conical probe that had directionsensing pressure ports. Although this

survey was made at PR5 = 1.69 (freestream Mach number, 0.9) because the

probe was not calibrated for supersonic

flow, the results are considered representative of the exit flow at higher

pressure ratios. Two survey traverses

10
I FLOW

1.0

!8

.7

-- O0

PARC3D

0 EXPERIMENTAL

0

/t)

MEASUREDON VENTRAL f

DUCTWALL JUST C) f

DOWNSTREAMOF .-

TAILPIPEOPENING-<..

"_0

(a) TAILPIPEWALL ON VENTRALSIDE.

0 0 0 0

.9

LOCATIONALONGTAILPIPEWALL

(b)TAILPIPEWALL OPPOSITEVENTRALOPENING.

FIG 14.- TAILPIPEWALL STATICPRESSURE.

at the exit plane were made: one near

an outboard edge of the nozzle

(Fig 15(a)), and the other near the

plane of symmetry (Fig 15(b)). Near

the outboard edge there was little or

no pressure loss except near the nozzle

forward lip. The flow direction was

outboard and forward in the aft part of

the nozzle and inboard and forward in

the front part. Near• the plane of symmetry the measured total pressure was

as much as 15 percent less than the

average station 5 pressure. The flow

directions were similar to the directions nearer the outboard edge, but the

flow angles were greater.

The results show that the flow leaving

the nozzle was still trying to "fill

in" the lower-density region at the

center of the upstream wall. Although

not measured, mirror images of these

patterns must exist on the other side

of the plane of symmetry. On that

basis the lateral velocity components

would cancel but cause a net thrust

loss, whereas the axial components

would add to cause a net thrust in the

reverse (downstream) direction. These

data, then, corroborate and explain the

axial force measured by the facility

load cells, previously shown in Fig 9.

An additional survey of the nozzle exit

flow was made with a total-pressure

rake. The data from this survey were

processed into a pressure contour map

at the exit plane, which is shown in

Fig 16 together with a similar map computed by the PARC3D code. 0nly one of

the symmetric halves is shown. Results

are given for PR5 = 3. The analytical

map always shows true total pressure,

but the experimental data are pitot

pressures, which are less than true

total pressure by the normalshock loss

at the tip in supersonic flow and by

measurement errors when the flow angle

relative to the tip is greater than

about 15 e. Nevertheless, the maps are

in good qualitative agreement: both

show strong flow in the rear part of

the nozzle and an oval-shaped region of

weaker flow in the front part.

CFD Modeling and Numerical Issues

Turbulence model - The Baldwin-Lomax

algebraic turbulence model was developed for two-dimensional, separated

flows. Extension of this model to

three-dimensional flows is difficult,

especially for cases having multiple

walls such as the rectangular ventral

duct. For the present problem the turbulent viscosities calculated for each

wall were weighted and averaged according to the distance from the point of

calculation to the relevant wall. No

adjustments to this approach were made

along walls in vortical regions. Inaccuracies involved in applying the

Baldwin-Lomax model in this manner may

have caused part of the differences

between the measured and predicted wall

static pressures.

A two-equation turbulence model, such

as the k-_ model (Ref i0), would handle

multiple walls in a more satisfactory

manner but would increase the computation time significantly_

Boundary-layer resolution in ventral

duc___!t- Resolution of boundary-layer

flow influenced CFD calculations

ii
N

IITRAVERSE I

PLANE OF_ I PATH --- _ tt _ -- IIOEFSETREAR EDGE

OF NOZZLE

FLOW

I STATION 5

I!I

I I ] I I I I

20--

-20 -- Z_ FLOW ANGLE IN PITCH PLANE

(POSITIVE FOR FLOW MOVING

TOWARD FRONT EDGE OF NOZZLE)

[] FLOW ANGLE IN YAW PLANE

-40 -- (POSITIVE FOR FLOW MOVING --

TOWARD OUTBOARD EDGE FOR /-FRONT

TRAVERSE PATH SHOWN) / EDGE

-60_-_ ....I....I I I I 1 1

0 I 2 3 h 5 6 7 0 I 2

DISTANCE FROM REAR EDGE OF NOZZLE, IN.

1 1 I 1 I t L J

0 4 8 12 16 20 0 4

DISTANCE FRO_ REAR EDGE OF NOZZLE, CM

(a) TRAVERSE PATH OFFSET 4.5 IN. (11.4 CM) FROM

PLAI_ OF SYMRETRY.

FIG 15. - VENTRAL NOZZLE EXIT FLOW CONDITIONS.

FRONT

EDGE-,

I l I I] 1

q 5 6 7

I I I I

8 12 1B 20

(b) TRAVERSE PATH OFFSET 1.5 iN. (3.8 CM) FROM

PLANE OF SYI_I_TRY.

RATIO OF TAILPIPE TO AMBIENT PRESSURE, PRG, 1.69.

OUTBOARD WALL

!:C!!;;!I'!ii!ii:::,

(a) EXPERI_NTAL PITOT PRESSURE.

FIG 16. - NOZZLE EXIT PLANE PRESSURE CONTOURS.

OUTBOARD WALL

F-MODEL .;:,.::: ::: • _.;: _;

_I PLANE :';:i-':)._i] :

SYMMETRY .j: -:::': : ....... _: _ :/_v\ I

(b) ANALYT]CAL TOTAL PRESSURE.

LIGHT SHADINGS INDICATE LOWER PRESSURES. DARK SHADINGS INDICATE HIGHER PRESSURES.

22
related to wall static pressure. For

the computational grids used in this

problem the y+ distance [iii] for

the first grid point of the ventral

duct wall was typically 15 for the fine

grid. This distance located the point

slightly outside the viscous sublayer

region of the boundary layer. In order

to fully define the boundary layer, the

first grid point should be at a y+ distance less than i0 from the wall surface (Ref Ii). However, decreasing the

first grid point y+ from 150 (for the

coarse grid) to 15 (for the fine grid)

did not change the calculated wall

static pressures significantly. A possible reason for this result is that

the ventral duct flow separated at the

sharp corners of the tailpipe opening

and was still turning in the duct.

Thus, because the boundary layer did

not build up in the usual way from viscosity and adverse pressure gradients,

the referenced y+ criterion may not be

applicable.

Conservation of mass - The inlet (station 5) and exit (station 6B) mass flow

rates computed by PARC3D with the fine

grid are shown in Fig 17 as a function

of the cumulative number of iterations.

As the solution progressed, the difference between the computed inlet and

exit flows diminished. One reason for

the flow difference is that the transfer of flow properties at the grid

block interface was not forced to be

conservative or characteristically correct. When flow convergence was

achieved, the residuals (differences in

flow properties between successive

iterations) had been reduced approximately three orders of magnitude. Further reduction seems to be limited by

the turbulence model used or possibly

by flow-field unsteadiness in the

vortices.

Low Mach number effects - The Mach number in the blocked tailpipe was low,

typically 0.i or less. As with other

compressible-flow codes the PARC3D code

encountered numerical problems in this

[iii] y+ = [(normal distance) (shear

stress/density)°'S]/kinematic

viscosity.

67

g 66

f

_ ss

_ 64

"TAILPIPE INLET

_3 / -NOZZLE EXIT

l

62 / t I 1 I

4 6 8 10 12x103

NUMBEROF ITERATIONS

FIG 17.- CALCULATEDMASS FLOWRATESAS FUNCTIONOF CUMULATIVE

NUMBEROF ITERATIONS.

region that led to convergence difficulties. Use of preconditioning

(Ref 12) could alleviate this problem.

Computational speed - The maximum

Courant-Friedrichs-Lewy (CFL) number

(Ref 13) that could be used for both

blocks to obtain a stable solution was

1.0 for the coarse grid and 0.5 for the

fine grid. (The CFL number limits the

iteration step size.) The computational speed for the coarse grid was

800 iterations per CPU hour on the Cray

X-MP. The fine-grid speed was 200

iterations per CPU hour on the Cray

Y-MP. The coarse-grid solution

required approximately 4000 iterations;

the fine-grid solution required 12 000

iterations.

Concluding Remarks

Experimental and analytical flow studies of the same generic model tailpipe

and ventral nozzle have been made. The

model was about one-third of full size,

and the end was closed to simulate a

blocked exhaust nozzle. Test data were

obtained up to a ratio of tailpipe to

ambient pressure of 4.0. The analytical work was done by using the PARC3D

computational fluid dynamics program to

predict the internal flow patterns and

overall ventral system performance.

The major results of these studies are

as follows:

(i) A flow-turning total-pressure loss

of about 5% percent was measured in the

model tested. This result is expected

to be typical of full-size systems of

similar geometry, but the loss should

be lower for smaller ventral nozzles or

13
systemswith entrance flow-turning

devices.

(2) Ventral nozzle flow and thrust

coefficients were normal considering

the measuredflow-turning loss.

(3) A reverse (directed downstream)

axial thrust componentwasmeasured,

although the ventral nozzle axis was

normal to the tailpipe centerline.

This indicates that the flow turned

morethan the intended 90°.

(4) Paint streak flow visualization

photographs and other data showed that

a low-density region of separated vortical flow occurred at the upstream

wall of the ventral duct. Flow was

strong in the downstream part of the

duct and tended to move toward the

upstream wall from both the side and

rear walls. This pattern persisted

through the nozzle exit and caused the

axial component measured by the thrust

system.

(5) The modeling technique and the

PARC3D computational code did a capable

job of analyzing internal flow patterns

and predicting system performance.

Solutions were obtained for two computational grid densities. Both gave

good results. The finer-grid solution

produced more detailed flow patterns

and predicted performance parameters,

such as thrust and flow coefficient,

within 1 percent of the measured

values.

3.

4.

5.

6.

8.

Raman, G., Zaman, K.B.M.Q., and Rice, E.J.,

"Initial Turbulence Effect on Jet Evolution With

and Without Tonal Excitation," AIAA Paper 87-

2725, Oct. 1987. (Also, NASA TM-100178.)

Pulliam, T.H. and Steger, _I.L., "Implicit FiniteDifference Simulations of Three Dimensional Compressible Flow, " AIAA Journal, Vol. 18, No. 2,

Feb. 1980, pp. 159-167.

Baldwin, B.S. and Lomax, H., "Thin-Layer

Approximation and Algebraic Model for Separated

Turbulent Flows," AIAA Paper 78-257, Jan. 1978.

Phares, W.J., Cooper, G.K., Jones, R.R., and Swafford,T.W., "Application of Computational Fluid

Dynamics to Test Facilityand Experiment

Design," AIAA Paper 86-1733, June 1986.

Stokes, M.L. and Kneile, K.L., "A search/

Interpolation Algorithm for CFD Analysis," presented at the World Congress on Computational

Mechanics, University of Texas, Austin, TX,

Sept. 1986.

Son], B.K., "Two- and Three-Dimensional Grid

Generation for Internal Flow Applications of Computational Fluid Dynamics," 7th Computational

Fluid Dynamics Conference, AIAA, 1985, pp. 351-

359.

9. Wlatka, P.P. and Baaing, P.J., "PLOT3D User's

Manual," NASA TM-101067, 1990.

10.

ll.

(6) Experience with the PARC3D code on

this problem indicates that computa- 12.

tional details could be improved further by using a two-equation turbulence

model (such as the k-_ model) and preconditioning in flow regions of low 13.

(<0.i) Mach number.

References

Wynosky, T.A. and Szyszko, C.J., "V/STOL

Deflector Aerodynamic Design Criteria - Geometric

Variations Effects," AIAA Paper 73-1181,

Nov. 1973.

1.

Wynosky, T.A., Streib,R.A., and Usab, W.J.,

"V/STOL DeflectorConcepts," AIAA Paper 74-

1168,Oct. 1974.

Harlow, F.H. and Nakayama, P.I.,"Transport of

Turbulence Energy Decay Rate," Los Alarnos ScientificLaboratory Report LA-3854, Los Alamos,

NM, 1968.

2.

Kuethe, A.M. and Chow, C.Y., Foundations of

Aerodynamics: Bases of Aerodynamic Design,

4th ed.,John Wiley & Sons, Inc.,New York, 1986.

Choi, Y.H., "Computation of Low Mach Number

Compressible Flow," Ph.D. Thesis,The Pennsylvania StateUniversity,1989.

Courant, R., Friedrichs,K.O., and Lewy, H.,

"Uber die PartieUen Differenzenglelchungender

Mathematischen Physik," Mathematische Annalen,

Vol. 100, 1928,pp. 32-74. (Translated to: "On the

PartialDifferenceEquations of Mathematical Physics,"IBM Journal of Research and Development,

Vol. If,1967,pp. 215-243.)

14
Report Documentation Page National Aeronauticsand

SpaceAdministration

1. Report NO. 2. Government Accession No.

NASA TM-103170

4. Title and Subtitle

Experimental and Analytical Study of Close-Coupled Ventral Nozzles for

ASTOVL Aircraft

7. Author(s)

Jack G. McArdle and C. Frederic Smith

9. Performing Organization Name and Address

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

3. Recipient's Catalog No.

5. Report Date

6. Performing Organization Code

8. Performing Organization Report No.

E-5499

10. Work Unit No.

505-62-71

11. Contract or Grant No.

13. Type of Reportand PeriodCovered

Technical Memorandum

14. Sponsoring Agency Code

15. Supplementary Notes

Prepared for the International Powered Lift Conference sponsored by the Royal Aeronautical Society,

London, England, August 29-31, 1990.

16. Abstract

Flow in a generic ventral nozzle system was studied experimentally and analytically with a block version of the

PARC3D computational fluid dynamics program (a full Navier-Stokes equation solver) in order to evaluate the

program's ability to predict system performance and internal flow patterns. For the experimental work a onethird-size model tailpipe with a single large rectangular ventral nozzle mounted normal to the tailpipe axis was

tested with unheated air at steady-state pressure ratios up to 4.0. The end of the tailpipe was closed to simulate a

blocked exhaust nozzle. Measurements showed about 51⁄2:percent flow-turning loss, reasonable nozzle

performance coefficients, and a significant aftward axial component of thrust due to flow turning more than 90*.

Flow behavior into and through the ventral duct is discussed and illustrated with paint streak flow visualization

photographs. For the analytical work the same ventral system configuration was modeled with two computational

_d'_:to evaluate the effect of grid density. Both grids gave good results. The finer-grid solution produced moredetailed flow patterns and predicted performance parameters, such as thrust and discharge coefficient, within

1 percent of the measured values. PARC3D flow visualization images are shown for comparison with the paint

streak photographs. Modeling and computational issues encountered in the analytical work are discussed.

17. Key Words (Suggested by Author(s))

VSTOL; STOVL; Ventral nozzles; CFD; Flow

visualization; Exhaust systems; Exhaust nozzles;

PARC3D; Navier-Stokes solutions

18. Distribution Statement

Unclassified- Unlimited

Subject Category 07

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of pages

Unclassified Unclassified 16

NASAFORMlS2eocrae *For sale by the National Technical Information Service, Springfield, Virginia 22161

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