NASA Technical Memorandum 103123 

AIAA-90-1899 

Analysis of Internal Flow 

in a Ventral Nozzle for 

STOVL Aircraft 

C. Frederic Smith 

Sverdrup Technology, Inc. 

Lewis Research Center Group 

Brook Park, Ohio 

and 

Jack G. McArdle 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 

Prepared for the 

26th Joint Propulsion Conference 

cosponsored by the AIAA, SAE, ASME, and ASEE 

Orlando, Florida, July 16-18, 1990 

Nl\SI\ 

NASA Technical Memorandum 103123 

AIAA-90-1899 

Analysis of Internal Flow 

in a Ventral Nozzle for 

STOVL Aircraft 

C. Frederic Smith 

Sverdrup Technology, Inc. 

Lewis Research Center Group 

Brook Park, Ohio 

and 

Jack G. McArdle 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 

Prepared for the 

26th Joint Propulsion Conference 

cosponsored by the AIAA, SAE, ASME, and ASEE 

Orlando, Florida, July 16-18, 1990 

Nl\SI\ 
ANALYSIS OF INTERNAL FLOW IN A VENTRAL 

NOZZLE FOR STOVL AIRCRAFT 

C. Frederic Smith" 

Sverdrup Technology , Inc. 

Lewis Research Center Group 

Brook Park, Ohio 44142 

and 

Jack G. McArdle t 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 44135 

Summary 

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, hover, 

and horizontal fught. To accomplish this, many designs include 

a ventral nozzle to provide part of the vertical thrust required. 

Understanding and predicting the internal aerodynamic flow 

caused by a single exhaust duct opening are highly desirable 

in assessing this concept. This paper presents a numerical 

simulation of a ventral nozzle and compares the results with 

experimental data. Comparisons include visualizations of the 

flow along the ventral duct walls and in the tailpipe plane of 

symmetry. Performance calculations are also compared with 

measured values. 

Introduction 

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, hover, 

and supersonic horizontal flight. In the lift mode the rear jet 

nozzle is blocked, and valves are opened to duct engine exhaust 

gases to two or more thrusters that are directed downward . 

In many proposed STOVL configurations, one of the lift 

• Assistant Supervisor, Propulsion AerOdynami cs Section. 

t Aerospace Research Engineer. 

thrusters is a ventral nozzle which draws mixed core and fan 

gases from the engine tailpipe through a valve and opening, 

and has no inJet turning vanes. Close coupling between the 

tailpipe and the ventral nozzle is necessary because the valve 

and nozzle must be mounted wholly within the fuselage . The 

ventral nozzle also may swivel to provide trim and pitch 

control. An example of a general configuration of a ventral 

nozzle in an aircraft is shown in figure 1. 

The concept of a ventral nozzle was explored in the early 

1970's. The advantage of this approach for vertical thrust 

generation is that the nozzle can be positioned such that the 

total lift vector is located very near the aircraft's center of 

gravity. Early work on STOVL deflectors is reported by 

Wynosky et aJ. (refs. 1 and 2), whose experiments examined 

the internal flow field and performance of a turbofan engine 

and ventral nozzle test model configuration. A major concern 

discussed in reference 1 is the upstream static-pressure 

gradients created at the turbine exit by turning the flow 90° 

into the ventral duct. These gradients, or flow distortions, 

could induce flow stability problems in the fan or turbine. The 

experiments indicate that this distortion diminishes rapidly with 

increasing distance from the ventral nozzle opening. 

The objective of this current effort is to demonstrate the 

capability to numerically model the internal flow in a generic 

tailpipe and ventral nozzle configuration. To accomplish this goal, 

numerical predictions were made by solving the full threedimensional Navier-Stokes equations with an algebraic 

turbulence model. These results were compared with experimental data for one configuration. The comparisons include 

ANALYSIS OF INTERNAL FLOW IN A VENTRAL 

NOZZLE FOR STOVL AIRCRAFT 

C. Frederic Smith" 

Sverdrup Technology , Inc. 

Lewis Research Center Group 

Brook Park, Ohio 44142 

and 

Jack G. McArdle t 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 44135 

Summary 

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, hover, 

and horizontal fught. To accomplish this, many designs include 

a ventral nozzle to provide part of the vertical thrust required. 

Understanding and predicting the internal aerodynamic flow 

caused by a single exhaust duct opening are highly desirable 

in assessing this concept. This paper presents a numerical 

simulation of a ventral nozzle and compares the results with 

experimental data. Comparisons include visualizations of the 

flow along the ventral duct walls and in the tailpipe plane of 

symmetry. Performance calculations are also compared with 

measured values. 

Introduction 

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, hover, 

and supersonic horizontal flight. In the lift mode the rear jet 

nozzle is blocked, and valves are opened to duct engine exhaust 

gases to two or more thrusters that are directed downward . 

In many proposed STOVL configurations, one of the lift 

• Assistant Supervisor, Propulsion AerOdynami cs Section. 

t Aerospace Research Engineer. 

thrusters is a ventral nozzle which draws mixed core and fan 

gases from the engine tailpipe through a valve and opening, 

and has no inJet turning vanes. Close coupling between the 

tailpipe and the ventral nozzle is necessary because the valve 

and nozzle must be mounted wholly within the fuselage . The 

ventral nozzle also may swivel to provide trim and pitch 

control. An example of a general configuration of a ventral 

nozzle in an aircraft is shown in figure 1. 

The concept of a ventral nozzle was explored in the early 

1970's. The advantage of this approach for vertical thrust 

generation is that the nozzle can be positioned such that the 

total lift vector is located very near the aircraft's center of 

gravity. Early work on STOVL deflectors is reported by 

Wynosky et aJ. (refs. 1 and 2), whose experiments examined 

the internal flow field and performance of a turbofan engine 

and ventral nozzle test model configuration. A major concern 

discussed in reference 1 is the upstream static-pressure 

gradients created at the turbine exit by turning the flow 90° 

into the ventral duct. These gradients, or flow distortions, 

could induce flow stability problems in the fan or turbine. The 

experiments indicate that this distortion diminishes rapidly with 

increasing distance from the ventral nozzle opening. 

The objective of this current effort is to demonstrate the 

capability to numerically model the internal flow in a generic 

tailpipe and ventral nozzle configuration. To accomplish this goal, 

numerical predictions were made by solving the full threedimensional Navier-Stokes equations with an algebraic 

turbulence model. These results were compared with experimental data for one configuration. The comparisons include 
paint fl ow visualizations, total pressure su rveys, wall static 

pressures, and performance data. 

Experiment Description 

A generic model tailpipe (about one-third full scale) with 

a single, large ventral nozzle was built and tested at NASA 

Lewis Research Center's Powered Lift Facility (PLF) to study 

the internal flow in this type of configuration. The facility was 

supplied with unheated air. A photograph of the experimental 

model is shown in figure 2. (Note that the model was tested 

upside down to simplify the experiment.) Thi model is based 

on a General Electric preliminary design proposed for the 

General Dynamics E-7D aircraft (ref. 3). The geometry of 

the experiment, although similar to that of the E-7D, was 

simplified to accommodate a straightforward grid-generation 

procedure for a numerical model of the flow field . However, 

the simplified experimental model still contains the essential 

flow physics of the E-7D. 

A schematic of the experimental model is shown in figure 3. 

As flow entered the tailpipe, it passed through a screen and 

a wire tripped the boundary layer. Then the flow was forced 

to turn 90° to exit through the ventral nozzle. A blind flange 

at the end of the tailpipe simulated a blocked exhaust nozzle 

for the vertical flight mode. Measurements included flow, 

forces, and internal pressures. In addition, a thin plate was 

mounted along the model tailpipe centerline at the ventral 

opening to visualize the flow with dabs of oily paint. The paint 

ran along streamlines when airflow was established, and the 

resulting streaks provided a picture of the flow paths into 

the ventral opening. Paint flow visualizations were also performed along the ventral duct walls without the tailpipe flow 

visualization plate present. A detailed pitot-pressure survey 

was conducted at the nozzle exit (station 6b , fig. 3). Static 

pressures were also measured along the ta ilpipe and ventral 

duct/nozzle walls. For the data in this report, the ratio of 

tailpipe total pressure to ambient pressure was 2.96. 

Numerical Analysis 

PARC3D Code 

The analytical performance of this configuration was studied 

with the PARC3D code and the results were compared with 

experimental data to establish confidence in the ability of this 

code to reasonably predict the flow field in a tailpipe and 

ventral nozzle configuration. The PARC3D code was developed 

at the ASA Ames Research Center (ref. 4) for analyzing 

external flows. This code solves the compressible, threedimensional Navier-Stokes equations in generalized curvilinear 

coordinates with an algebraic turbulence model. It was later 

extended to analyze internal flows at the Arnold Engineering 

Development Center (ref. 5). PARC3D uses a Baldwin-Lomax 

turbulence model (ref. 6), and it solves the equations of motion 

2 

by using the Beam-Warming approximate factorization scheme. 

Pulliam's scalar pentadiagonal transformation is applied to 

uncouple the equations of motion . This approach provides an 

efficient steady-state solver. 

A block version of the PARC3D code was used. One advantage of this approach is that the computational grid is divided 

into two or more grid blocks to simplify the modeling of a 

complex geometry. Trilinear interpolation (ref. 7) is used to 

transfer information at grid block interfaces. Another advantage 

of using a block solver is that the majority of the information 

that needs to be stored in core memory is limited to the current 

block being analyzed. Therefore a problem involving a very 

large number of grid points may be solved with a computer 

that has less core memory than would be required to solve 

the problem by using two or more grid blocks. 

Grid and Boundary Conditions 

A "wire mesh " diagram of the tailpipe and ventral nozzle 

configuration is shown in figure 4. Since the experimental 

hardware had a geometric plane of symmetry, only half of 

the configuration was modeled for the computations. Since 

the tailpipe and ventral nozzle were at right angles to each 

other, it would have been very difficult to use one contiguous 

grid to model these components adequately. Therefore, the 

grid model was divided into two blocks: the tailpipe and the 

ventral nozzle. The tailpipe was modeled with an O-grid , and 

the ventral duct/nozzle was modeled with an H-grid. The 

O-grid consisted of concentric circles that were parallel to 

the tailpipe surface and radial lines that were perpendicular 

to this surface. The H-grid provided lines that were perpendicular and parallel to the ventral duct/nozzle walls. This 

approach provided body-conforming grids for each block. 

These grids were generated algebraically by the INGRID3D 

code (ref. 8). The grids were stretched by hyperbolic stretching 

functions and were packed near the walls and near the center 

of the O-grid. Two grids were used for this study to evaluate 

the effects of the grid size on the predicted flow field. The 

first grid contained 31 875 grid points in each block (51 by 

25 by 25 points in the streamwise and orthogonal directions, 

respectively). This grid was limited by computer disk space 

available on the NASA Lewis Cray X-MP computer, but it 

was deemed adequate to evaluate the applicability of the 

PARC3D code to this problem. In order to provide an adequate 

number of points in the core flow region of the nozzle, grid 

packing near the nozzle surface was limited and the first grid 

point for the ventral nozzle was placed 0.1 in . from the surface. 

Three views of this grid are shown in figure 5. The results 

obtained with this grid were reasonable in comparison with 

the data. 

A finer grid was generated for the NAS Cray Y - MP computer to explore the effects of grid density on the numerical 

solution. The finer grid contained 262 701 points per block 

( 101 by 51 by 51 points) . Three views similar to the coarse 

grid are shown in figure 6 for comparison with figure 5. For 

---- --- ------ ------------------------------------

paint fl ow visualizations, total pressure su rveys, wall static 

pressures, and performance data. 

Experiment Description 

A generic model tailpipe (about one-third full scale) with 

a single, large ventral nozzle was built and tested at NASA 

Lewis Research Center's Powered Lift Facility (PLF) to study 

the internal flow in this type of configuration. The facility was 

supplied with unheated air. A photograph of the experimental 

model is shown in figure 2. (Note that the model was tested 

upside down to simplify the experiment.) Thi model is based 

on a General Electric preliminary design proposed for the 

General Dynamics E-7D aircraft (ref. 3). The geometry of 

the experiment, although similar to that of the E-7D, was 

simplified to accommodate a straightforward grid-generation 

procedure for a numerical model of the flow field . However, 

the simplified experimental model still contains the essential 

flow physics of the E-7D. 

A schematic of the experimental model is shown in figure 3. 

As flow entered the tailpipe, it passed through a screen and 

a wire tripped the boundary layer. Then the flow was forced 

to turn 90° to exit through the ventral nozzle. A blind flange 

at the end of the tailpipe simulated a blocked exhaust nozzle 

for the vertical flight mode. Measurements included flow, 

forces, and internal pressures. In addition, a thin plate was 

mounted along the model tailpipe centerline at the ventral 

opening to visualize the flow with dabs of oily paint. The paint 

ran along streamlines when airflow was established, and the 

resulting streaks provided a picture of the flow paths into 

the ventral opening. Paint flow visualizations were also performed along the ventral duct walls without the tailpipe flow 

visualization plate present. A detailed pitot-pressure survey 

was conducted at the nozzle exit (station 6b , fig. 3). Static 

pressures were also measured along the ta ilpipe and ventral 

duct/nozzle walls. For the data in this report, the ratio of 

tailpipe total pressure to ambient pressure was 2.96. 

Numerical Analysis 

PARC3D Code 

The analytical performance of this configuration was studied 

with the PARC3D code and the results were compared with 

experimental data to establish confidence in the ability of this 

code to reasonably predict the flow field in a tailpipe and 

ventral nozzle configuration. The PARC3D code was developed 

at the ASA Ames Research Center (ref. 4) for analyzing 

external flows. This code solves the compressible, threedimensional Navier-Stokes equations in generalized curvilinear 

coordinates with an algebraic turbulence model. It was later 

extended to analyze internal flows at the Arnold Engineering 

Development Center (ref. 5). PARC3D uses a Baldwin-Lomax 

turbulence model (ref. 6), and it solves the equations of motion 

2 

by using the Beam-Warming approximate factorization scheme. 

Pulliam's scalar pentadiagonal transformation is applied to 

uncouple the equations of motion . This approach provides an 

efficient steady-state solver. 

A block version of the PARC3D code was used. One advantage of this approach is that the computational grid is divided 

into two or more grid blocks to simplify the modeling of a 

complex geometry. Trilinear interpolation (ref. 7) is used to 

transfer information at grid block interfaces. Another advantage 

of using a block solver is that the majority of the information 

that needs to be stored in core memory is limited to the current 

block being analyzed. Therefore a problem involving a very 

large number of grid points may be solved with a computer 

that has less core memory than would be required to solve 

the problem by using two or more grid blocks. 

Grid and Boundary Conditions 

A "wire mesh " diagram of the tailpipe and ventral nozzle 

configuration is shown in figure 4. Since the experimental 

hardware had a geometric plane of symmetry, only half of 

the configuration was modeled for the computations. Since 

the tailpipe and ventral nozzle were at right angles to each 

other, it would have been very difficult to use one contiguous 

grid to model these components adequately. Therefore, the 

grid model was divided into two blocks: the tailpipe and the 

ventral nozzle. The tailpipe was modeled with an O-grid , and 

the ventral duct/nozzle was modeled with an H-grid. The 

O-grid consisted of concentric circles that were parallel to 

the tailpipe surface and radial lines that were perpendicular 

to this surface. The H-grid provided lines that were perpendicular and parallel to the ventral duct/nozzle walls. This 

approach provided body-conforming grids for each block. 

These grids were generated algebraically by the INGRID3D 

code (ref. 8). The grids were stretched by hyperbolic stretching 

functions and were packed near the walls and near the center 

of the O-grid. Two grids were used for this study to evaluate 

the effects of the grid size on the predicted flow field. The 

first grid contained 31 875 grid points in each block (51 by 

25 by 25 points in the streamwise and orthogonal directions, 

respectively). This grid was limited by computer disk space 

available on the NASA Lewis Cray X-MP computer, but it 

was deemed adequate to evaluate the applicability of the 

PARC3D code to this problem. In order to provide an adequate 

number of points in the core flow region of the nozzle, grid 

packing near the nozzle surface was limited and the first grid 

point for the ventral nozzle was placed 0.1 in . from the surface. 

Three views of this grid are shown in figure 5. The results 

obtained with this grid were reasonable in comparison with 

the data. 

A finer grid was generated for the NAS Cray Y - MP computer to explore the effects of grid density on the numerical 

solution. The finer grid contained 262 701 points per block 

( 101 by 51 by 51 points) . Three views similar to the coarse 

grid are shown in figure 6 for comparison with figure 5. For 

---- --- ------ ------------------------------------
the ventral nozzle, the first grid point was 0.01 in. from the 

surface. For the tailpipe in both grids, the first grid point was 

0.01 in. from the surface. Comparisons of figures 5 and 6 

indicate much denser packing of the grid near the walls and 

tailpipe center when the fi ne grid was used. The grid was 

packed near the center to minimize the effects of the " pole " 

boundary condition, which is discussed later. 

The boundary conditions for this problem are shown in 

figure 7. Because it was desirable to keep the size of the 

computational grid as small as possible, the flow field outside 

of the nozzle was not analyzed. Although proper modeling of 

the nozzle exit flow conditions is very important, the application of boundary conditions at the nozzle throat is difficult 

because of the transonic nature of the flow. This problem was 

solved by adding a fictitious diverging section to the ventral 

nozzle to provide supersonic flow at the exit of the 

computational domain. This added section allowed the flow 

properties to be extrapolated downstream of the area of interest and modeled the nozzle throat as if the external plume were 

included. Previous studies have shown this to be a good 

approach to modeling a three-dimensional converging nozzle. 

Another problem occurred at the center of the tailpipe 

O-grid: the radial grid lines became coincident at the center, 

causing singularity problems in the calculation of the coordinate transformation derivatives (metrics). To circumvent this 

problem, a so-caHed pole boundary condition was employed 

at the center of the O-grid. The flow properties were averaged 

around the adjacent grid line, and these average values were 

applied to all points along the innermost grid line, which has 

a radius of approximately 1 percent of the tailpipe radius. 

Results and Discussion 

The discussion in thjs section is divided into several topics. 

The comparisons of predicted particle trajectories with paint 

flow visualizations are discussed first. This is followed 

by comparisons of measured and predicted total and static 

pressures, along with performance parameters. The last topic 

concerns numerical modeling and computational aspects of the 

solutions obtained. The program PLOT3D (ref. 9) was used 

to generate the graphical presentation of the numerical results. 

Flow Visualization 

The following discussion compares paint flow visualizations 

and predicted particle trajectories obtained with the coarse and 

fme grids. These trajectories represent flow paths that massless 

particles would foHow if they were released into the flow field. 

In figure 8, the tailpipe plane of symmetry paint flow results 

are compared with the fine and coarse grid solutions. As can 

be seen in the paint flows, a large vortex exists in the blocked 

tailpipe region. The predictions for the fine grid solution 

indicate several smaller vortices that are not present in the 

coarse grid solution. These small vortices are not apparent in 

the paint flow studies. This may be due to the paint not being 

able to fl ow within such small regions. The predicted vortex 

center is farther from the tailpipe centerline than the paint flows 

indicate. One source for this discrepancy may be the no-slip 

surface on the plate used in the visualization. The computed 

particle trajectories are along an inviscid plane of symmetry. 

The discontinuities in the particle paths at the tailpipe centerline 

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 upstream ventral duct wall paint flows are shown in 

figure 9 along with the predicted particle trajectories. (The 

nomenclature for the ventral duct walls is presented in fig. 6.) 

The two solutions are similar to each other and compare well 

with the experimental data . As can be seen from the flow 

patterns, the ventral duct vortex pulls fluid towards the plane 

of symmetry from the side wall of the duct. The two counterrotating vortices can be seen in the paint streaks and the particle 

trajectories along the duct wall moving towards each other. 

The downstream ventral-duct wall flow visualization is shown 

in figure 10 along with predictions. In this region, the finer grid 

provides much better agreement with the experiment than the 

coarse grid does. In this figure, the flow is split into two distinct directions. ear the duct inlet, the flow tends to move back 

towards the duct inlet, which represents a region of separated 

flow. Farther along the duct wall, the flow moves away from 

the side ventral duct wall and towards the nozzle exit. 

The side duct wall paint flows are shown in figure 11 . Both 

accompanying predictions agree well qualitatively with the 

experimental data. In this figure the flow tends to move from 

the downstream wall towards the upstream duct wall. This 

movement was due to the presence of the nozzle vortices. 

Total and Static Pressures 

The pitot pressures measured in the nozzle exit plane and 

the total pressure fi eld predicted by the P ARC3D code are 

shown in figure 12. An oval-shaped region of low pressure 

is observed near the upstream nozzle wall in both the measured 

data and the numerical results. The solutions obtained with 

the coarse and fi ne grids provide simjiar results. The fine grid 

results indicate a stronger vortex than the coarse grid solution 

indicates. Measured and predicted flow-angle surveys show 

that the nozzle flow from the downstream and side walls is 

moving towards this region. The predicted contours of local 

total to upstream total pressure ratios greater than 1.01 are 

restricted to regions very close to the ventral duct walls. This 

is not apparent in the color contour plots shown. 

These flow characteristics are exhibited well in the predicted 

particle trajectories shown in figure 13. A major feature 

shown in this figure is the large vortical region near the 

upstream wall of the ventral duct. The lower pressure in 

this region drew flow from the downstream and side ventral 

duct walls. The vortex persisted through the nozzle exit. This 

vortical region is shown in the paint flow vi ualization of 

3 

the ventral nozzle, the first grid point was 0.01 in. from the 

surface. For the tailpipe in both grids, the first grid point was 

0.01 in. from the surface. Comparisons of figures 5 and 6 

indicate much denser packing of the grid near the walls and 

tailpipe center when the fi ne grid was used. The grid was 

packed near the center to minimize the effects of the " pole " 

boundary condition, which is discussed later. 

The boundary conditions for this problem are shown in 

figure 7. Because it was desirable to keep the size of the 

computational grid as small as possible, the flow field outside 

of the nozzle was not analyzed. Although proper modeling of 

the nozzle exit flow conditions is very important, the application of boundary conditions at the nozzle throat is difficult 

because of the transonic nature of the flow. This problem was 

solved by adding a fictitious diverging section to the ventral 

nozzle to provide supersonic flow at the exit of the 

computational domain. This added section allowed the flow 

properties to be extrapolated downstream of the area of interest and modeled the nozzle throat as if the external plume were 

included. Previous studies have shown this to be a good 

approach to modeling a three-dimensional converging nozzle. 

Another problem occurred at the center of the tailpipe 

O-grid: the radial grid lines became coincident at the center, 

causing singularity problems in the calculation of the coordinate transformation derivatives (metrics). To circumvent this 

problem, a so-caHed pole boundary condition was employed 

at the center of the O-grid. The flow properties were averaged 

around the adjacent grid line, and these average values were 

applied to all points along the innermost grid line, which has 

a radius of approximately 1 percent of the tailpipe radius. 

Results and Discussion 

The discussion in thjs section is divided into several topics. 

The comparisons of predicted particle trajectories with paint 

flow visualizations are discussed first. This is followed 

by comparisons of measured and predicted total and static 

pressures, along with performance parameters. The last topic 

concerns numerical modeling and computational aspects of the 

solutions obtained. The program PLOT3D (ref. 9) was used 

to generate the graphical presentation of the numerical results. 

Flow Visualization 

The following discussion compares paint flow visualizations 

and predicted particle trajectories obtained with the coarse and 

fme grids. These trajectories represent flow paths that massless 

particles would foHow if they were released into the flow field. 

In figure 8, the tailpipe plane of symmetry paint flow results 

are compared with the fine and coarse grid solutions. As can 

be seen in the paint flows, a large vortex exists in the blocked 

tailpipe region. The predictions for the fine grid solution 

indicate several smaller vortices that are not present in the 

coarse grid solution. These small vortices are not apparent in 

the paint flow studies. This may be due to the paint not being 

able to fl ow within such small regions. The predicted vortex 

center is farther from the tailpipe centerline than the paint flows 

indicate. One source for this discrepancy may be the no-slip 

surface on the plate used in the visualization. The computed 

particle trajectories are along an inviscid plane of symmetry. 

The discontinuities in the particle paths at the tailpipe centerline 

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 upstream ventral duct wall paint flows are shown in 

figure 9 along with the predicted particle trajectories. (The 

nomenclature for the ventral duct walls is presented in fig. 6.) 

The two solutions are similar to each other and compare well 

with the experimental data . As can be seen from the flow 

patterns, the ventral duct vortex pulls fluid towards the plane 

of symmetry from the side wall of the duct. The two counterrotating vortices can be seen in the paint streaks and the particle 

trajectories along the duct wall moving towards each other. 

The downstream ventral-duct wall flow visualization is shown 

in figure 10 along with predictions. In this region, the finer grid 

provides much better agreement with the experiment than the 

coarse grid does. In this figure, the flow is split into two distinct directions. ear the duct inlet, the flow tends to move back 

towards the duct inlet, which represents a region of separated 

flow. Farther along the duct wall, the flow moves away from 

the side ventral duct wall and towards the nozzle exit. 

The side duct wall paint flows are shown in figure 11 . Both 

accompanying predictions agree well qualitatively with the 

experimental data. In this figure the flow tends to move from 

the downstream wall towards the upstream duct wall. This 

movement was due to the presence of the nozzle vortices. 

Total and Static Pressures 

The pitot pressures measured in the nozzle exit plane and 

the total pressure fi eld predicted by the P ARC3D code are 

shown in figure 12. An oval-shaped region of low pressure 

is observed near the upstream nozzle wall in both the measured 

data and the numerical results. The solutions obtained with 

the coarse and fi ne grids provide simjiar results. The fine grid 

results indicate a stronger vortex than the coarse grid solution 

indicates. Measured and predicted flow-angle surveys show 

that the nozzle flow from the downstream and side walls is 

moving towards this region. The predicted contours of local 

total to upstream total pressure ratios greater than 1.01 are 

restricted to regions very close to the ventral duct walls. This 

is not apparent in the color contour plots shown. 

These flow characteristics are exhibited well in the predicted 

particle trajectories shown in figure 13. A major feature 

shown in this figure is the large vortical region near the 

upstream wall of the ventral duct. The lower pressure in 

this region drew flow from the downstream and side ventral 

duct walls. The vortex persisted through the nozzle exit. This 

vortical region is shown in the paint flow vi ualization of 

3 
the upstream ventral duct wall and in the total-pressure surveys 

at the nozzle exit. The low-pressure region caused the jet 

to exjt the nozzle with an upstream-facing velocity component, 

thus producing a signjficant downstream component of thrust 

whjch was measured to be approximately 10 percent of the 

vertical thrust. The predicted down tream component, 

determined from the calculated average flow angles, is 

approximately 8 percent of the vertical thrust. As was shown 

in the experimental flow visualizations, there were counterrotating vortices present in the ventral duct. These vortices 

are consistent with the symmetric nature of the test hardware . 

The second vortex is assumed to be present in the calculations 

by the plane of symmetry boundary condition. The calcul ated 

tailpipe wall static pressure are within 2 percent of the 

measured data, and the predicted wall static pressures near 

the nozzle exit are within approximately 3 percent of the 

data. 

Figure 14(a) shows how the flow turning into the ventral 

duct affects the static-pressure gradients in the tajlpipe. The 

static-pressure field distortion decreases rapidly as the distance 

from the ventral nozzle opening increases. This trend implies 

that the ventral duct may not affect turbine or fan flow stability 

unless the turbine or fan is placed very close to the nozzle 

operung. This result is consistent with trends reported in 

references 1 and 2. The static pressures in the ventra l duct 

and nozzle are shown in figure 14(b). Large gradients in the 

static-pressure field due to the vortex present in the nozzle 

operung are observed. The flow field djstortion is also apparent 

in the Mach number contours shown in the tajlpipe and ventral 

nozzle hown in figure 15. In figure IS(a), a rapid acceleration 

of the flow around the upstream ventral duct corner can be 

seen. In fig ure lS(b), the presence of the vortex is indicated 

by the large gradients in the somewhat circular Mach number 

contours that are centered near the upstream plane of symmetry 

corner of the ventral duct and nozzle. 

Performance 

The predicted thrust and mass flow obtained from the coarse 

and fine grid solutions are compared with experimental data 

in table I. The ± numbers indicate the amount of oscillation 

in the solution. As can be seen, the fi ner grid exhibited a 

reduced amplitude. These oscillations may be due to the 

inherent unsteadiness of flows containing large vortical 

regions. The thrust predictions are in good agreement with 

the measured thrust, and the predicted mass flow compa res 

well with the experimental data. 

Modeling and Numerical Issues 

Static Pressure 

The discrepancies between the measured and predicted static 

pressures may be due, in part, to the turbulence model used. 

The Baldwin-Lomax turbulence model was developed for two4 

TA BLE I.- VE TRAL NOZZLE PERFORMANCE 

Perfo rmance indicators Numerical Ex pe rimental 

results results 

Coarse grid Fine grid 

Tailpipe 10 amb ient 2.97 2.96 2.96 

pressure ratio' P5/Pamb 

Nozzle inlet 10 tai lpipe .932 .935 .945 

pressure ratio . aP61P5 

Nozz le discharge coeffi c ient .88 .91 .91 

based on aP5 

Nozzle thrust coeffIcient .96 .97 .97 

based on aP5 

Corrected now rate. b Ibm/sec 22.67 ±0.2 23.32±0.04 23 .04 

Corrected thrusl.c Ibf 870±9 905 ±2 896 

Corrected axial thrust,C Ibf -88 -72 -73 

3Whcrc Ps = Iolal pressure al Sial Ion 5. P6=lotaJ pressure 31 Station 6. and Pamb = t0l31 amblcnt pressure. 

bel/hO)/o: where ti,=mass flow raiC. 8=(total tcmperarure)/5 18.76 R. b=(tOlal pressure at station 5),' 1~ 696 ps13 . 

C FIb: where F = thrust. 

dimensional, separated flows, and the extension of this model 

to three-dimensional flows is difficult, especially in cases 

where there are multiple walls (such as the ventral 

duct/nozzle). The turbulent viscosities calculated fo r each wall 

were weighted and averaged on the basis of the distance from 

the assoc iated wall to the point of calculation. In addition , the 

ventral duct flow contained a large vorti cal region that may 

not be accounted for adequately in the model. A two-equation 

turbulence model , such as a k-E model, would handle the 

multiple walls implicitly for the turbulence model equations. 

Boundary Layer Resolution 

Another consideration is that the boundary layer may not 

be resolved well enough. The y + distance for the fi rst grid 

point off the ventral duct wall was typically IS for the fine 

grid , where 

+ normal distance (shear stress/density)o.5 

y = kjnematic viscosity 

This distance is slightly outside of the viscous sublayer region 

of the boundary layer. For a fully defined boundary layer, the 

fir t grid point in the flow field should be located at a y + 

distance of less than 10 from the surface (ref. 10). However, 

decreasing the y + for the first grid point off the wall from 

ISO (for the coa rse grid) to IS (for the fine grid) did not 

significantly change the results for the wall static-pressure 

calculations. This may be due to the fact that the separation 

was generated by the sharp corner at the ventral duct opening 

and not by adverse pressure gradients. 

the upstream ventral duct wall and in the total-pressure surveys 

at the nozzle exit. The low-pressure region caused the jet 

to exjt the nozzle with an upstream-facing velocity component, 

thus producing a signjficant downstream component of thrust 

whjch was measured to be approximately 10 percent of the 

vertical thrust. The predicted down tream component, 

determined from the calculated average flow angles, is 

approximately 8 percent of the vertical thrust. As was shown 

in the experimental flow visualizations, there were counterrotating vortices present in the ventral duct. These vortices 

are consistent with the symmetric nature of the test hardware . 

The second vortex is assumed to be present in the calculations 

by the plane of symmetry boundary condition. The calcul ated 

tailpipe wall static pressure are within 2 percent of the 

measured data, and the predicted wall static pressures near 

the nozzle exit are within approximately 3 percent of the 

data. 

Figure 14(a) shows how the flow turning into the ventral 

duct affects the static-pressure gradients in the tajlpipe. The 

static-pressure field distortion decreases rapidly as the distance 

from the ventral nozzle opening increases. This trend implies 

that the ventral duct may not affect turbine or fan flow stability 

unless the turbine or fan is placed very close to the nozzle 

operung. This result is consistent with trends reported in 

references 1 and 2. The static pressures in the ventra l duct 

and nozzle are shown in figure 14(b). Large gradients in the 

static-pressure field due to the vortex present in the nozzle 

operung are observed. The flow field djstortion is also apparent 

in the Mach number contours shown in the tajlpipe and ventral 

nozzle hown in figure 15. In figure IS(a), a rapid acceleration 

of the flow around the upstream ventral duct corner can be 

seen. In fig ure lS(b), the presence of the vortex is indicated 

by the large gradients in the somewhat circular Mach number 

contours that are centered near the upstream plane of symmetry 

corner of the ventral duct and nozzle. 

Performance 

The predicted thrust and mass flow obtained from the coarse 

and fine grid solutions are compared with experimental data 

in table I. The ± numbers indicate the amount of oscillation 

in the solution. As can be seen, the fi ner grid exhibited a 

reduced amplitude. These oscillations may be due to the 

inherent unsteadiness of flows containing large vortical 

regions. The thrust predictions are in good agreement with 

the measured thrust, and the predicted mass flow compa res 

well with the experimental data. 

Modeling and Numerical Issues 

Static Pressure 

The discrepancies between the measured and predicted static 

pressures may be due, in part, to the turbulence model used. 

The Baldwin-Lomax turbulence model was developed for two4 

TA BLE I.- VE TRAL NOZZLE PERFORMANCE 

Perfo rmance indicators Numerical Ex pe rimental 

results results 

Coarse grid Fine grid 

Tailpipe 10 amb ient 2.97 2.96 2.96 

pressure ratio' P5/Pamb 

Nozzle inlet 10 tai lpipe .932 .935 .945 

pressure ratio . aP61P5 

Nozz le discharge coeffi c ient .88 .91 .91 

based on aP5 

Nozzle thrust coeffIcient .96 .97 .97 

based on aP5 

Corrected now rate. b Ibm/sec 22.67 ±0.2 23.32±0.04 23 .04 

Corrected thrusl.c Ibf 870±9 905 ±2 896 

Corrected axial thrust,C Ibf -88 -72 -73 

3Whcrc Ps = Iolal pressure al Sial Ion 5. P6=lotaJ pressure 31 Station 6. and Pamb = t0l31 amblcnt pressure. 

bel/hO)/o: where ti,=mass flow raiC. 8=(total tcmperarure)/5 18.76 R. b=(tOlal pressure at station 5),' 1~ 696 ps13 . 

C FIb: where F = thrust. 

dimensional, separated flows, and the extension of this model 

to three-dimensional flows is difficult, especially in cases 

where there are multiple walls (such as the ventral 

duct/nozzle). The turbulent viscosities calculated fo r each wall 

were weighted and averaged on the basis of the distance from 

the assoc iated wall to the point of calculation. In addition , the 

ventral duct flow contained a large vorti cal region that may 

not be accounted for adequately in the model. A two-equation 

turbulence model , such as a k-E model, would handle the 

multiple walls implicitly for the turbulence model equations. 

Boundary Layer Resolution 

Another consideration is that the boundary layer may not 

be resolved well enough. The y + distance for the fi rst grid 

point off the ventral duct wall was typically IS for the fine 

grid , where 

+ normal distance (shear stress/density)o.5 

y = kjnematic viscosity 

This distance is slightly outside of the viscous sublayer region 

of the boundary layer. For a fully defined boundary layer, the 

fir t grid point in the flow field should be located at a y + 

distance of less than 10 from the surface (ref. 10). However, 

decreasing the y + for the first grid point off the wall from 

ISO (for the coa rse grid) to IS (for the fine grid) did not 

significantly change the results for the wall static-pressure 

calculations. This may be due to the fact that the separation 

was generated by the sharp corner at the ventral duct opening 

and not by adverse pressure gradients. 
I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

Conservation of Mass 

The predicted inlet and exit mass flow variations for the fine 

grid solution as a function of number of time steps are shown 

in figure 16. The differences may be partiall y attributed to 

the fact that the transfer of flow properties at the interface 

between the grid blocks is not forced to be conservative or 

characteristicaJly correct. As the solution is iterated, the 

amplitude of the osciJlations and the discrepancies in the inlet 

and exit mass flow rates diminish. The residuals (differences 

in flow properties between successive iterations) were reduced 

approximately three orders of magnitude. Difficulties encountered in further reducing the size of the residuals may be 

attributable to the turbulence model or possible flow field 

unsteadiness. 

Low Mach Number Effects 

Another contributing factor to the convergence difficulties 

is the low Mach number flow in the blocked region of the 

tailpipe. Typical Mach numbers are less than 0.1 in this region; 

such low Mach levels are a classic source of numerical 

problems for compressible flow codes. Preconditioning could 

improve this situation (ref. 11). 

Computational Issues 

The maximum Courant-Friedrichs-Lewy (CFL) number 

(ref. 12) 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 is the time step limiter). The computational 

speed for the coarse grid was 800 iterations per CPU hr on 

the Cray X-MP. The fine-grid speed was 200 iterations per 

CPU hr on the Cray Y-MP. The coarse-grid solution required 

approximately 4000 iterations, whereas the fine-grid solution 

required 12 000 iterations. 

Conclusions and Recommendations 

To understand and predict the internal flow in a tailpipe and 

ventral nozzle configuration, a comparison was made between 

the numerical results obtained from the PARC3D code and 

experimental data for this model. 

The experimental and numerical resuJts showed that the fl ow 

field within a tailpipe and ventral nozzle configuration is 

extremely complex. The results of this study demonstrate that 

the PARC3D code can predict realistic flow fields and performance for such a configuration. The numerical solutions 

are somewhat grid dependent. Several flow field details exhibited in the fine-grid results are not present in the coarse-grid 

solutions, although reasonable predictions of the major flow 

field phenomena and nozzle performance parameters were 

made using the coarse grid. The analysis reconfirmed previous 

experimental results that the flow turning into the ventral duct 

opening has a limited effect on the upstream flow distortion 

in the tailpipe for the present configuration. 

One major step for improving the numerical model would 

be to apply a two-equation turbulence model in place of the 

algebraic model currently used. The two-equation model would 

handle multiple walls in a more straightforward manner. A 

k-E (two-equation) turbulence model is being incorporated into 

the P ARC3D code. The computational time required to use 

such a turbulence model will increase considerably. 

The flow field has large regions of very low speed flow 

which can impede convergence for compressible flow solvers. 

The inclusion of preconditioning may alleviate these problems. 

Acknowledgments 

This work was managed by Peter G. Batterton under NASA 

Contract NAS3- 25266. Part of the analysis was performed 

utilizing the Numerical Aerodynamic Simulation (NAS) 

fac ility located at the NASA Ames Research Center. 

References 

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

Design Criteria-Geometric Variations Effects," AlAA Paper 73- 11 81, 

Nov. 1973. 

2. Wynosky, T.A., Streib, R.A. , and Usab, W.J. , " V/STOL Deflector 

Concepts," AIAA Paper 74-1168, Oct. 1974. 

3. Meyer, B.E. , Re, R.J. , and Romine, B.M., " A Static Investigation of 

Several STOVL Exhaust System Concepts," AIAA Paper 89-2928, July 

1989. 

4. Pulliam, T.H., and Steger, J.L. , " Implicit Finite-Difference Simulations 

of Three Dimensional Compressible Flow," AIM Journal, Vol. 18, 

No.2, Feb. 1980, pp. 159- 167. 

5. Cooper, G.K., Jones, R.R., Phares, W.J., and Swafford , T.W. , 

" Application of Computational Fluid Dynamics to Test Facility and 

Experiment Design," A1AA Paper 86-1733, June 1986. 

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

Algebraic Turbu lence Model for Separated Turbulent Flows, " AIAA 

Paper 78-257, Jan. 1978. 

7. 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. 

8. Soni , B.K. , " Two- and Three-Dimensional Grid Generation fo r Internal 

Flow Applications of Computational Fluid Dynamics, " AIAA Paper 

85- 1526, July 1985. 

9. Wlatka , P.P., and Buning, P.G. , "PLOT3D User's Manual ," NASA 

TM-IOI067, 1990. 

10. Kuethe, A.M. , and Chow, C. Y.: Foundations of Aerodynamics, Bases 

of Aerodynamic Design, 4th ed. , John Wiley and Sons, Inc. , New York, 

1986. 

I I. Choi, Y.H., " Computation of Low Mach Number Compressible Flow," 

Ph .D. Thesis, The Pennsylvania State University , 1989. 

12. Courant, R., Friedrichs, K.O., and Lewy , H. , "Uber die Partiellen 

Differenzengleichungen der Mathematischen Physik, " Mathematische 

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

Difference Equations of Mathematical Physics," IBM Journal of 

Research and Development, Vol. II, 1967, pp . 215-243. 

5 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

Conservation of Mass 

The predicted inlet and exit mass flow variations for the fine 

grid solution as a function of number of time steps are shown 

in figure 16. The differences may be partiall y attributed to 

the fact that the transfer of flow properties at the interface 

between the grid blocks is not forced to be conservative or 

characteristicaJly correct. As the solution is iterated, the 

amplitude of the osciJlations and the discrepancies in the inlet 

and exit mass flow rates diminish. The residuals (differences 

in flow properties between successive iterations) were reduced 

approximately three orders of magnitude. Difficulties encountered in further reducing the size of the residuals may be 

attributable to the turbulence model or possible flow field 

unsteadiness. 

Low Mach Number Effects 

Another contributing factor to the convergence difficulties 

is the low Mach number flow in the blocked region of the 

tailpipe. Typical Mach numbers are less than 0.1 in this region; 

such low Mach levels are a classic source of numerical 

problems for compressible flow codes. Preconditioning could 

improve this situation (ref. 11). 

Computational Issues 

The maximum Courant-Friedrichs-Lewy (CFL) number 

(ref. 12) 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 is the time step limiter). The computational 

speed for the coarse grid was 800 iterations per CPU hr on 

the Cray X-MP. The fine-grid speed was 200 iterations per 

CPU hr on the Cray Y-MP. The coarse-grid solution required 

approximately 4000 iterations, whereas the fine-grid solution 

required 12 000 iterations. 

Conclusions and Recommendations 

To understand and predict the internal flow in a tailpipe and 

ventral nozzle configuration, a comparison was made between 

the numerical results obtained from the PARC3D code and 

experimental data for this model. 

The experimental and numerical resuJts showed that the fl ow 

field within a tailpipe and ventral nozzle configuration is 

extremely complex. The results of this study demonstrate that 

the PARC3D code can predict realistic flow fields and performance for such a configuration. The numerical solutions 

are somewhat grid dependent. Several flow field details exhibited in the fine-grid results are not present in the coarse-grid 

solutions, although reasonable predictions of the major flow 

field phenomena and nozzle performance parameters were 

made using the coarse grid. The analysis reconfirmed previous 

experimental results that the flow turning into the ventral duct 

opening has a limited effect on the upstream flow distortion 

in the tailpipe for the present configuration. 

One major step for improving the numerical model would 

be to apply a two-equation turbulence model in place of the 

algebraic model currently used. The two-equation model would 

handle multiple walls in a more straightforward manner. A 

k-E (two-equation) turbulence model is being incorporated into 

the P ARC3D code. The computational time required to use 

such a turbulence model will increase considerably. 

The flow field has large regions of very low speed flow 

which can impede convergence for compressible flow solvers. 

The inclusion of preconditioning may alleviate these problems. 

Acknowledgments 

This work was managed by Peter G. Batterton under NASA 

Contract NAS3- 25266. Part of the analysis was performed 

utilizing the Numerical Aerodynamic Simulation (NAS) 

fac ility located at the NASA Ames Research Center. 

References 

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

Design Criteria-Geometric Variations Effects," AlAA Paper 73- 11 81, 

Nov. 1973. 

2. Wynosky, T.A., Streib, R.A. , and Usab, W.J. , " V/STOL Deflector 

Concepts," AIAA Paper 74-1168, Oct. 1974. 

3. Meyer, B.E. , Re, R.J. , and Romine, B.M., " A Static Investigation of 

Several STOVL Exhaust System Concepts," AIAA Paper 89-2928, July 

1989. 

4. Pulliam, T.H., and Steger, J.L. , " Implicit Finite-Difference Simulations 

of Three Dimensional Compressible Flow," AIM Journal, Vol. 18, 

No.2, Feb. 1980, pp. 159- 167. 

5. Cooper, G.K., Jones, R.R., Phares, W.J., and Swafford , T.W. , 

" Application of Computational Fluid Dynamics to Test Facility and 

Experiment Design," A1AA Paper 86-1733, June 1986. 

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

Algebraic Turbu lence Model for Separated Turbulent Flows, " AIAA 

Paper 78-257, Jan. 1978. 

7. 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. 

8. Soni , B.K. , " Two- and Three-Dimensional Grid Generation fo r Internal 

Flow Applications of Computational Fluid Dynamics, " AIAA Paper 

85- 1526, July 1985. 

9. Wlatka , P.P., and Buning, P.G. , "PLOT3D User's Manual ," NASA 

TM-IOI067, 1990. 

10. Kuethe, A.M. , and Chow, C. Y.: Foundations of Aerodynamics, Bases 

of Aerodynamic Design, 4th ed. , John Wiley and Sons, Inc. , New York, 

1986. 

I I. Choi, Y.H., " Computation of Low Mach Number Compressible Flow," 

Ph .D. Thesis, The Pennsylvania State University , 1989. 

12. Courant, R., Friedrichs, K.O., and Lewy , H. , "Uber die Partiellen 

Differenzengleichungen der Mathematischen Physik, " Mathematische 

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

Difference Equations of Mathematical Physics," IBM Journal of 

Research and Development, Vol. II, 1967, pp . 215-243. 

5 
6 

BLOCKED END 

'\ EJECTOR 

\ 

L VENTRAL 

VALVE AND 

NOZZLE 

Figure I. - Ventral nozzle concept. 

VENTRAL NOZZLE 

\ 

\ 

EXHAUST \ 

NOZZLE ~ 

1\ I LOW INLET -I 

I 

I 

I 

C-89- 10105 

Figure 2.- Tailpipe and ventral nozzle experimental hardwa re. ( ote that the model was tested upside down to simplify the experiment.) 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

6 I 

BLOCKED END 

'\ EJECTOR 

\ 

L VENTRAL 

VALVE AND 

NOZZLE 

Figure I. - Ventral nozzle concept. 

VENTRAL NOZZLE 

\ 

\ 

EXHAUST \ 

NOZZLE ~ 

1\ I LOW INLET -I 

I 

I 

I 

C-89- 10105 

Figure 2.- Tailpipe and ventral nozzle experimental hardwa re. ( ote that the model was tested upside down to simplify the experiment.) 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 
/ 

6.~-IN. X l1.7- IN . ( 16.2-CM X 29.7-CM) 

\ VENTRAL NOZZLE \ 

\ 

I 

13 1 N . 

-(33 CM)-

/ 

/ 

/ 

12 IN. 

(30 CM) 

,-- BLI ND 

I FLANGE 

STATION 6 (2~ 

TOTAL PRESSURES)---

L OPT I ONAl FLOW 

r BOUNDARY 

/ lAYER TRIP I 

/ ,SCREEN AND FLOW 

/ I STRA I GHTENER 

\ ~=::::, ::;~ 

\ 

\ 

\ 

V I SUALI ZA TI ON 

PLATE 

I OPTIONAL FLOW~ 

l PLATE 

7 I N.(68 . 6 CM) . : VI SUALI ZATION 

MODEL TA I LP IPE LENGTH. 8.7A IN. ~ CM) - ---TRANSITION SECTION----.! 

SECTION A-A 

/ 

/ 

L BLOCKED 

EN D 

Figure 3.- Schematic of experimental model. 

, , 

FLOW 

' ,- TAILPIPE 

Figure 4.-"Wire mesh" model of tailpipe and ventral nozzle. 

• FLOW 

FAC I LITY 

MOUNTI NG 

FLANGE 

7 

/ 

6.~-IN. X l1.7- IN . ( 16.2-CM X 29.7-CM) 

\ VENTRAL NOZZLE \ 

\ 

I 

13 1 N . 

-(33 CM)-

/ 

/ 

/ 

12 IN. 

(30 CM) 

,-- BLI ND 

I FLANGE 

STATION 6 (2~ 

TOTAL PRESSURES)---

L OPT I ONAl FLOW 

r BOUNDARY 

/ lAYER TRIP I 

/ ,SCREEN AND FLOW 

/ I STRA I GHTENER 

\ ~=::::, ::;~ 

\ 

\ 

\ 

V I SUALI ZA TI ON 

PLATE 

I OPTIONAL FLOW~ 

l PLATE 

7 I N.(68 . 6 CM) . : VI SUALI ZATION 

MODEL TA I LP IPE LENGTH. 8.7A IN. ~ CM) - ---TRANSITION SECTION----.! 

SECTION A-A 

/ 

/ 

L BLOCKED 

EN D 

Figure 3.- Schematic of experimental model. 

, , 

FLOW 

' ,- TAILPIPE 

Figure 4.-"Wire mesh" model of tailpipe and ventral nozzle. 

• FLOW 

FAC I LITY 

MOUNTI NG 

FLANGE 

7 
8 

/ 

PLANE OF / 

SYMMETRY -< , 

, 

" 

~ VENTRAL DUCT 

/ AND NOZILE / 

WALL --

BLOCKED ___ _E"--1111 • FLOW 

( bl 

,,-- VENTRAL DUCT 

AN D NOZZLE 

r- S I DE WALL / 

TAILPIPE 

'- TAILP IPE 

DOWN STREAM 

WALL ', , 

(a l 

(a) Plane of symmetry (xz-plane). 

(b) End view (yz-plane). 

(c) Nozzle exit (xy-plane). 

Figure 5.-Coarse grid . 

r- PLANE OF SYMMETRY / 

"-'-- S I DE WALL 

(el 

/ 

r- UPSTREAM 

WALL 

8 

/ 

PLANE OF / 

SYMMETRY -< , 

, 

" 

~ VENTRAL DUCT 

/ AND NOZILE / 

WALL --

BLOCKED ___ _E"--1111 • FLOW 

( bl 

,,-- VENTRAL DUCT 

AN D NOZZLE 

r- S I DE WALL / 

TAILPIPE 

'- TAILP IPE 

DOWN STREAM 

WALL ', , 

(a l 

(a) Plane of symmetry (xz-plane). 

(b) End view (yz-plane). 

(c) Nozzle exit (xy-plane). 

Figure 5.-Coarse grid . 

r- PLANE OF SYMMETRY / 

"-'-- S I DE WALL 

(el 

/ 

r- UPSTREAM 

WALL 
PLANE OF 

SYMMETRY 

BLOCKED 

END--

(b) 

..-~ VENTRAL DUCT 

AND NOZZLE 

SIDE WALL 

TAILPIPE 

,,- VENTRAL DUCT 

(a) 

DOWNSTRE AM 

WALL - --

/ 

• FLOW 

TAILPIPE 

,,- PLANE OF SYMMETRY 

/--

-

(a) Plane of symmetry (xz-plane). 

(b) End view (yz-plane). 

(e) Nozzle exit (xy-plane). 

Figure 6. - Fine grid . 

, 

'-SI DE WALL 

(c) 

UPSTREAM 

WALL 

9 

PLANE OF 

SYMMETRY 

BLOCKED 

END--

(b) 

..-~ VENTRAL DUCT 

AND NOZZLE 

SIDE WALL 

TAILPIPE 

,,- VENTRAL DUCT 

(a) 

DOWNSTRE AM 

WALL - --

/ 

• FLOW 

TAILPIPE 

,,- PLANE OF SYMMETRY 

/--

-

(a) Plane of symmetry (xz-plane). 

(b) End view (yz-plane). 

(e) Nozzle exit (xy-plane). 

Figure 6. - Fine grid . 

, 

'-SI DE WALL 

(c) 

UPSTREAM 

WALL 

9 
10 

EXTRAPOLATE 

(MACH NUMBER > 1) _/ 

POLE 

;' 

;' 

/' 

----.: .... 

~F ICTITIOU S 

DIVERGI NG 

SECTION 

', ~ TRUE NOZZLE 

EX IT PLANE 

/ - NO SLIP, 

ADIABATI C 

/~ I NTERPOLATE 

~ NO SLIP , 

AD I ABATIC 

t PLANE OF SYMMETRY 

Figure 7.- Boundary conditions. Total pressure and total temperature were 

specified for upstream portion of ta il pipe; adiabatic conditions and no slip 

were specified for downstream portion of tai lpipe. 

---- - ------

10 

EXTRAPOLATE 

(MACH NUMBER > 1) _/ 

POLE 

;' 

;' 

/' 

----.: .... 

~F ICTITIOU S 

DIVERGI NG 

SECTION 

', ~ TRUE NOZZLE 

EX IT PLANE 

/ - NO SLIP, 

ADIABATI C 

/~ I NTERPOLATE 

~ NO SLIP , 

AD I ABATIC 

t PLANE OF SYMMETRY 

Figure 7.- Boundary conditions. Total pressure and total temperature were 

specified for upstream portion of ta il pipe; adiabatic conditions and no slip 

were specified for downstream portion of tai lpipe. 

---- - ------
(c) 

(d) 

(a) Coarse grid . (b) Fine grid 

(c) Experiment. 

(d) Prediction. 

Figure 8.-Tai lpipe plane of symmetry fl ow visualization. 

--- VENTRAL 

DUCT 

C-89- 6731 

....... FLOW 

II 

(c) 

(d) 

(a) Coarse grid . (b) Fine grid 

(c) Experiment. 

(d) Prediction. 

Figure 8.-Tai lpipe plane of symmetry fl ow visualization. 

--- VENTRAL 

DUCT 

C-89- 6731 

....... FLOW 

II 
(a) Coa rse grid . (b) Fine grid . 

(c) Experiment. (Nozzle not present.) 

(d) Prediction. (Nozzle not present ; solution refl ected about plane of symmetry.) 

Figure 9 .-Upstream ventral duct and nozzle wall fl ow visualizati on. 

12 

(a) Coa rse grid . (b) Fine grid . 

(c) Experiment. (Nozzle not present.) 

(d) Prediction. (Nozzle not present ; solution refl ected about plane of symmetry.) 

Figure 9 .-Upstream ventral duct and nozzle wall fl ow visualizati on. 

12 
REVERSE 

PLANE OF 

SYMMETRY 

(a) Coar e grid. (b) Fine grid. 

(c) Experiment. (Nozzle not present.) 

(d) Prediction. ( ozzle not pre ent; solution reflected about plane of symmetry.) 

Figure IO.- Downstream ventral duct and nozzle wall flow visualization . 

13 

REVERSE 

PLANE OF 

SYMMETRY 

(a) Coar e grid. (b) Fine grid. 

(c) Experiment. (Nozzle not present.) 

(d) Prediction. ( ozzle not pre ent; solution reflected about plane of symmetry.) 

Figure IO.- Downstream ventral duct and nozzle wall flow visualization . 

13 
(el 

14 

UPSTREAM 

WA LL 

(a) Coarse grid . (b) Fine grid . 

(e) Experiment. (Nozzle not present.) 

(d) Prediction. ( ozzle not present.) 

Figure I I.- Side ventral duct and nozzle wa ll flow visuali zation. 

(el 

14 

UPSTREAM 

WA LL 

(a) Coarse grid . (b) Fine grid . 

(e) Experiment. (Nozzle not present.) 

(d) Prediction. ( ozzle not present.) 

Figure I I.- Side ventral duct and nozzle wa ll flow visuali zation. 
CONTDUR LEVEL:! 

n.QUlIOtJ 

fI.U7f1ll1l 

11 . 1'1 .... 1 

.... lIillI.1 

11.1«11 •• 1 

11.' 11 111.1 

11 . '111 11 . ' 

11 . !II>lJlIll 

11 . 9111.1' 

I.'I"U' 

UPSTREAM 

WAL L ___ 

COf'lTIIUW l(U(lS 

lI .n .. II"1 

II . (nn .. 1 

U. ·('9I""1 

11.11111.1 

1I . 1f3I1I1I1 

11 . 11'''''''' 11.'1111111 1 

".·Hllliol 

II ' II . "lIlIa l 

,I. ~I!"."I ' . II IIJI1lI 

1. tt: .. I.1 

•. tr"" . 11 1 

CONlOUM l(UClS 

II ,GSOIIIl 

O.G71lOIl 

0.&9I1UII 

1I . f1UIII 

1I .""'''"l 11. · ........ 111 

U.""UUII 

0.(90IlU 

1I.lIlIft"1 

01. '11'"'" 

11.'I:lI".' 

II II" I 

U. !lttJll' 

fI ~"II.1 

1. 101 .... 

1. ItoUII' 

\ 

\ 

/ - SIDE WALL 

~PLANE OF SYMMETRY 

( a) 

( b) 

( c) 

(a) Measured pitot pressure. 

(b) Coarse grid total pressure. 

(e) Fine grid total pressure. 

r DOWNSTftEAM 

I WAL L 

Figure 12.-Nozzle exit plane pitot and total pressures. (The contou rs represent levels of local total to upstream total pressure ratios.) 

15 

CONTDUR LEVEL:! 

n.QUlIOtJ 

fI.U7f1ll1l 

11 . 1'1 .... 1 

.... lIillI.1 

11.1«11 •• 1 

11.' 11 111.1 

11 . '111 11 . ' 

11 . !II>lJlIll 

11 . 9111.1' 

I.'I"U' 

UPSTREAM 

WAL L ___ 

COf'lTIIUW l(U(lS 

lI .n .. II"1 

II . (nn .. 1 

U. ·('9I""1 

11.11111.1 

1I . 1f3I1I1I1 

11 . 11'''''''' 11.'1111111 1 

".·Hllliol 

II ' II . "lIlIa l 

,I. ~I!"."I ' . II IIJI1lI 

1. tt: .. I.1 

•. tr"" . 11 1 

CONlOUM l(UClS 

II ,GSOIIIl 

O.G71lOIl 

0.&9I1UII 

1I . f1UIII 

1I .""'''"l 11. · ........ 111 

U.""UUII 

0.(90IlU 

1I.lIlIft"1 

01. '11'"'" 

11.'I:lI".' 

II II" I 

U. !lttJll' 

fI ~"II.1 

1. 101 .... 

1. ItoUII' 

\ 

\ 

/ - SIDE WALL 

~PLANE OF SYMMETRY 

( a) 

( b) 

( c) 

(a) Measured pitot pressure. 

(b) Coarse grid total pressure. 

(e) Fine grid total pressure. 

r DOWNSTftEAM 

I WAL L 

Figure 12.-Nozzle exit plane pitot and total pressures. (The contou rs represent levels of local total to upstream total pressure ratios.) 

15 
Figure l3.-Predicted particle trajectory paths. 

16 

Figure l3.-Predicted particle trajectory paths. 

16 
CONTOUR L UELS 

0 . 44000 

0 . 46000 

0 . 48000 

0 .50000 

0.57000 

0.54000 

0.56000 

J lilt J 

(J 6 on 

0.66000 

0.60000 

0.700011 

0.77000 

CO TOUR LEUElS 

0.44000 

0.46000 

0 . 48000 

0.50000 

0 .52000 

0.5'1000 

J .1 JI J 

n.h70no 

0 . 6·1000 

0.66000 

0.60000 

FLOW 

' FLOW 

(a) Tailpipe. 

(b) Ventral duci and nozzle. 

Figure 14.-Predicted static pressure. (Contours represent levels of normalized local static pressure.) 

17 

CONTOUR L UELS 

0 . 44000 

0 . 46000 

0 . 48000 

0 .50000 

0.57000 

0.54000 

0.56000 

J lilt J 

(J 6 on 

0.66000 

0.60000 

0.700011 

0.77000 

CO TOUR LEUElS 

0.44000 

0.46000 

0 . 48000 

0.50000 

0 .52000 

0.5'1000 

J .1 JI J 

n.h70no 

0 . 6·1000 

0.66000 

0.60000 

FLOW 

' FLOW 

(a) Tailpipe. 

(b) Ventral duci and nozzle. 

Figure 14.-Predicted static pressure. (Contours represent levels of normalized local static pressure.) 

17 
18 

CONTOUR LEUELS 

0 . 00000 

0 . 05000 

0 . 10000 

0 . 15 000 

0 .20000 

0 .25000 

0.30000 

aU11 

~rnoo 

0 .55000 

0 . 60000 

0 . 6 5000 

0 . 70000 

CONTOUR LEUELS 

0 . 000 00 

0 . 0 000 

0 . 10000 

0. 1 00 

O. ?OOOO 

0 .250 00 

0.30000 

0 .3500U 

I. ttlllUU 

. no 0.6000n 

0 . 65000 

0 .70000 

0. 75000 

0 . 00000 

( b) 

' FLOW 

Ca) Tailpipe. 

(b) Ventral duct and nozzle. 

Figure IS .-Predicted Mach number. 

18 

CONTOUR LEUELS 

0 . 00000 

0 . 05000 

0 . 10000 

0 . 15 000 

0 .20000 

0 .25000 

0.30000 

aU11 

~rnoo 

0 .55000 

0 . 60000 

0 . 6 5000 

0 . 70000 

CONTOUR LEUELS 

0 . 000 00 

0 . 0 000 

0 . 10000 

0. 1 00 

O. ?OOOO 

0 .250 00 

0.30000 

0 .3500U 

I. ttlllUU 

. no 0.6000n 

0 . 65000 

0 .70000 

0. 75000 

0 . 00000 

( b) 

' FLOW 

Ca) Tailpipe. 

(b) Ventral duct and nozzle. 

Figure IS .-Predicted Mach number. 
67 

~ 66 

'" "-E 

E 65 

.z 

a 

-' 64 u.. 

(/) 

(/) 

~ 63 

62 

4 6 8 

ITERATIONS 

Figure 16.-CaIculated mass flow rates as function of number of time steps. 

19 

67 

~ 66 

'" "-E 

E 65 

.z 

a 

-' 64 u.. 

(/) 

(/) 

~ 63 

62 

4 6 8 

ITERATIONS 

Figure 16.-CaIculated mass flow rates as function of number of time steps. 

19 
NI\SI\ 

National Report Documentation Page Aeronautics and 

Space Administration 

1. Report No. NASA TM- 103123 2. Government Accession No. 3. Recipient's Catalog No. 

AIAA -90- 1899 

4. Title and Subtitle 5. Report Date 

Analysis of Internal Flow in a Ventral Nozzle for STOVL Aircraft 

6. Performing Organization Code 

7. Author(s) 8. Performing Organization Report No. 

e. Frederic Smjth and Jack G. McArdle E-545 I 

10. Work Unit No. 

505-62-71 

9. Performing Organization Name and Address 

11 . Contract or Grant No. 

National Aeronautics and Space Admjnistration 

Lewis Research Center 

Cleveland, Ohio 44135-3191 13. Type of Report and Period Covered 

12. Sponsoring Agency Name and Address Technical Memorandum 

National Aeronautics and Space Administration 14. Sponsoring Agency Code 

Washington, D.e. 20546- 0001 

15. Supplementary Notes 

Prepared for the 26th Joint Propulsion Conference cosponsored by the AIAA, SAE, ASME, and ASEE, Orlando, 

Florida , July 16-18, 1990. e. Frederic Smith , Sverdrup Technology, Inc., Lewis Research Center Group, 

2001 Aerospace Parkway , Brook Park, Ohio 44142. Jack G. McArdle, NASA Lewis Research Center. 

16. Abstract 

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, hover, and horizontal flight. To accomplish this, many 

designs include a ventral nozzle to provide part of the vertical thrust required . Understanding and predicting the 

internal aerodynamjc flow caused by a single exhaust duct opening are highly desirable in assessing this concept. 

This paper presents a numerical simulation of a ventral nozzle and compares the results with experimental data. 

Comparisons include visualizations of the flow along the ventral duct walls and in the tailpipe plane of symmetry . 

Performance calculations are also compared with measured values. 

17. Key Words (Suggested by Author(s)) 18. Distribution Statement 

STOVL; Ventral nozzle; CFD, PARC3D; Flow Unclassified - Unlimited 

visualization; Navier-Stokes solutions; Exhaust nozzle Subj ect Category 07 

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21 . No. of pages 22. Price ' 

Unclassified Unclassified 22 A03 

NASA FORM 1626 OCT 86 *Fu sale by the National Technical Information Service, Springfield, Virginia 22161 

NI\SI\ 

National Report Documentation Page Aeronautics and 

Space Administration 

1. Report No. NASA TM- 103123 2. Government Accession No. 3. Recipient's Catalog No. 

AIAA -90- 1899 

4. Title and Subtitle 5. Report Date 

Analysis of Internal Flow in a Ventral Nozzle for STOVL Aircraft 

6. Performing Organization Code 

7. Author(s) 8. Performing Organization Report No. 

e. Frederic Smjth and Jack G. McArdle E-545 I 

10. Work Unit No. 

505-62-71 

9. Performing Organization Name and Address 

11 . Contract or Grant No. 

National Aeronautics and Space Admjnistration 

Lewis Research Center 

Cleveland, Ohio 44135-3191 13. Type of Report and Period Covered 

12. Sponsoring Agency Name and Address Technical Memorandum 

National Aeronautics and Space Administration 14. Sponsoring Agency Code 

Washington, D.e. 20546- 0001 

15. Supplementary Notes 

Prepared for the 26th Joint Propulsion Conference cosponsored by the AIAA, SAE, ASME, and ASEE, Orlando, 

Florida , July 16-18, 1990. e. Frederic Smith , Sverdrup Technology, Inc., Lewis Research Center Group, 

2001 Aerospace Parkway , Brook Park, Ohio 44142. Jack G. McArdle, NASA Lewis Research Center. 

16. Abstract 

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, hover, and horizontal flight. To accomplish this, many 

designs include a ventral nozzle to provide part of the vertical thrust required . Understanding and predicting the 

internal aerodynamjc flow caused by a single exhaust duct opening are highly desirable in assessing this concept. 

This paper presents a numerical simulation of a ventral nozzle and compares the results with experimental data. 

Comparisons include visualizations of the flow along the ventral duct walls and in the tailpipe plane of symmetry . 

Performance calculations are also compared with measured values. 

17. Key Words (Suggested by Author(s)) 18. Distribution Statement 

STOVL; Ventral nozzle; CFD, PARC3D; Flow Unclassified - Unlimited 

visualization; Navier-Stokes solutions; Exhaust nozzle Subj ect Category 07 

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21 . No. of pages 22. Price ' 

Unclassified Unclassified 22 A03 

NASA FORM 1626 OCT 86 *Fu sale by the National Technical Information Service, Springfield, Virginia 22161