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UNIVERSITY OF SOUTHERN CALIFORNIA

Reduction of Distortion in Bifurcated Intake Systems Utilizing Boundary Layer Ingestion

Sina Golshany Dec. 3, 2008

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Abstract: This report intends to present a solution to the problem of distortion created inside an aircraft’s bifurcated intake system. Utilizing a low intensity suction tool, boundary layers generated on the inner duct walls are removed. Therefore, a significant improvement in the distortion levels of a conventional bifurcated intake system could be achieved with minimum power extraction from the jet engine. This result may have significant design implications, allowing a larger number of jet engines to be considered for bifurcated intake system integration.

1. Introduction:

B

ifurcated intake systems are just as significant today as they were when Saab and Dassault first introduced them to the aviation industry over fifty years ago. From a

historical standpoint, many aircraft, mainly in the military domain, have utilized the bifurcated intake system. Having a single buried jet engine within the fuselage demands a different approach to the problem of intake integration; interferences within the structure, internal components, and the cockpit restrict the design efforts to two conventional approaches. S-duct intakes, which could be seen in commercial aircrafts like the Boeing 727 and Lockheed L-1011, present weaknesses

that

have

been

observed during their long service period. Even at a slight positive attack angle the flow entering the S-duct intake is partially disturbed

Fig. 1) Shown here is an inboard profile of S-duct integration for a very small jet trainer. Notice that the by the fuselage, thus causing an un-steady center supplyofof air to the buried engine. In many cases, gravity of the aircraft is shifted forward due the restrictions caused by the design of the intake system.

such disturbances have led to undesired engine shut down.

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Figure 1 presents an example of this intake integration methodology applied to a very light jet trainer. The alternative approach is to utilize a bifurcated intake, which is comprised of two merging ducts. In contrast to the S-duct, this design is not sensitive to external disturbances caused by high angle of attack and side-slip angle. A bifurcated intake system, however, faces

Fig. a2) an isometric view of a typical S-duct intake internal ↑ significant Fig. 3) an isometric view of the bifurcated intake system designed for the experiments proposed in this document.→

disturbance* created by the formation of boundary layers along the duct walls. Given the aforementioned advantages and the long history of utilization, bifurcated intakes are considered to be the most feasible low-weight solution to the problem of intake integration in modern jet fighter design philosophy. From a historical standpoint, a large number of aircraft, mainly in the military domain, have utilized the bifurcated intake system. Some of the more famous aircraft that have used a bifurcated intake system includes, Douglas A-4 Skyhawk, Lockheed F-104 Starfighter, Republic F-105 Thunderchief, Convair F-106 Delta Dart, and Lockheed U-2.

Also two of the most recent fighter designs, Lockheed F-35

Lighting, and Saab-39 Gripen utilize the bifurcated intake integration method. As it will be presented in this report, bifurcated intakes have an inherited inefficiency regarding the turbulent boundary layers in the merging flow. It has been the objective of this                                                             *

Disturbance is generally defined as the variation of the velocity and pressure profiles before the first stage of the engine. Velocity and pressure distributions are expected to be relatively independent of each other as it is established in the case of compressible flow. It could be caused by swirls and vortices in the air flow as well as undesired separated boundary layers inside the intake, or by external causes such as turbulent local flow adjacent to the intake.

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project to offer a simple solution to this inefficiency by introducing the concept of internal boundary layer removal, and to verify the practicality of this concept in a real world design scenario.

In order to achieve this objective, a series of physical experiments has been

conducted for which the process and results will be presented in this report. This project is initiated by the design problems encountered during the AIAA Individual Aircraft Design Competition. In order to present the complete thought process leading to the final design of apparatus, a brief review of the process of design of an aircraft intake is presented in the Appendix 1. Using the bifurcated intake designed for the design completion, a hypothesis was made about the merging boundary layers. It was expected that the boundary layers formed on the inner duct walls to merge and create a dorsal layer of slow fluid upon exiting the intake system. In order to validate this hypothesis in the designed bifurcated intake system, a transient simulation was performed using ANSYS CFD analysis module. Fluid properties for 35000 ft. were applied to the model boundaries following the atmospheric model presented by Roskam.

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The model was solved for fluid velocity and pressure in order to obtain the velocity and pressure profiles.

Fig. 4) Velocity Contour, V=100 m/sec

Fig. 6) Velocity Contour, V=180

Fig. 5) Pressure Contour, V=100 m/sec

Fig. 7) Pressure Contour, V=180 m/sec

Figure 4 and 5: The bifurcated intake is modeled using ANSYS CFD module (V=200 kts, Alt=35000 ft., Pref =21330 Pa). The turbulence is modeled using Shi-Zhu-Lumley in order to achieve a fast convergence while accounting for compressibility effects. As it can be seen from these figures, compressibility effect varies the density of the air and therefore causes the pressure on the intake plane not to be a quadratic function of the fluid velocity as it has been established for incompressible flow. Figure 6 and 7: The same analysis is performed with higher airspeed (V=350 kts.) and with the same fluid properties. As it can be seen from the velocity contours, the dorsal boundary layer from the duct walls is extended to the engine phase, causing a non-homogenous airspeed profile in cruise condition, the most important and prevalent condition.

As it can be seen from the CFD results presented in figures 4 and 6, the dorsal boundary layer is predicted to be present in a wild range of cruise velocities from 100 m/sec. to 180 m/sec.

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Reviewing the results, the average thickness of the boundary layer is measured to be 22 percent of the exit diameter of the intake (10.3 centimeter on average). From these results it is evident that the velocity and pressure drastically change in the dorsal region of the intake exit, leading to non-homogeneity of the flow as it enters the jet engine. Given that a significant distortion in the flow caused by the non-homogeneity of the velocity and pressure profile can inflict unwanted cyclic stress on fan blades and engine mechanism, it is considered beneficial to be able to reduce the thickness of the dorsal low-speed region, as well as the intensity of the velocity profile gradient in the dorsal region.

Fig. 8) Velocity Vectors, V=180 m/sec. A top view of the intake exit.

The goal of the project was Set to reduce the depression created in velocity profile in the dorsal region by 70 percent or more, and to reduce the thickness of the distorted dorsal region as much as possible. It can be argued that this design issue could be addressed by adding a low-pressure chamber at the intersection of both ducts in order to induce a weak suction on their surfaces, thereby reducing the boundary layer thickness along the inner walls and reducing the disturbance caused by the merging boundary layers in the airflow entering the jet engine.

1.2 Apparatus Design: In order to validate the feasibility of this solution, experiments are performed to determine the proper flow rate inside suction chamber device installed on the intersection of the two ducts. Openings on the duct walls are considered to allow the boundary layer to be removed due to the application of suction. In order to validate the performance of these slots

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in removing the boundary layer, a small-scale model of a bifurcated intake system, both including and dismissing the suction, is exposed to airflow inside the wind tunnel. The apparatus for this experiment is consisted of a small-scale model of the bifurcated intake system explained in the previous section to ensure geometric similarity of the airflow in the model with the real size intake. In order to achieve similitude between the airflow of the small scale model and the real size intake system, it was necessary to properly scale down the real-size intake system designed for the AIAA individual design competition.

Since the

phenomenon in study – the formation of boundary layers – is highly dependent on the characteristics of the air flow, it was also important to generate a turbulent flow inside the model. This implied that the Reynolds number should be at least in the order of 105 to cause the boundary layer to become turbulent. Also the apparatus is designed in order to have kinematic similarity to the real-size designed intake. This fact requires that the velocity profiles at the points of interest on the model (near the inner walls of the intake, where the reduction in the boundary layer thickness is to be observed), should be linearly proportional to the velocity profiles obtained from the CFD analysis of the real-size intake system. Therefore, the main design decisive factor for the sizing of apparatus was considered to be obtaining the similarity of the boundary layer velocity profile in the dorsal regions of the intake exit. Fox et al.† presents the disturbance thickness of the turbulent boundary layer formed along stationary smooth walls as:

 Turb  x Streamline

0.382 Re1x/ 5

(1)

                                                            †

R. W. Fox, P.J. Pritchard, A.T. McDonald, “Introduction to Fluid Mechanics 7TH Edition”, 2009, John Riley & Sons

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Where the parameter x streamline is the length measured along the inner curved wall of the intake system and the Reynolds number is calculated based on this length using (2):

U 1 x Streamline (2)  In this analysis, the effects of the engine on the intake are neglected, given that the method Re x 

used to design the original intake‡ neglects the effects of the engine on the intake system due to analytical limitations. To obtain an expression for the thickness of the boundary layer observed in the model, (2) is substituted into (1), together yielding:

 Turb  x Streamline

0.382  U 1 x Streamline   

1/ 5

  

(3)

Since Equation 3 only calculates the thickness of the boundary layer for one of the ducts, doubling this magnitude will yield the thickness of the boundary layer observed at the intersection of the two ducts: 2 Turb    Intake Exit Boundary Layer Thickness   (4) In order to validate the effects of the suction device, the horizontal velocity profile at the intake exit should be determined experimentally.

This can be accomplished by

performing velocity measurements at the exit cross section of the intake model using a pitot tube. Therefore it can be argued that the ratio of the thickness of the dorsal boundary layer to the diameter of the pitot tube could be used as a parameter for properly scaling down the model. After consulting with project advisors§, it was agreed upon that in order to make sufficient number of measurements to determine the velocity profile of the intake exit, the boundary layer thickness should be five to ten times the diameter of the Pitot tube. To perform the required analysis, (3) is solved for the length of the streamline.

                                                            ‡ §

J. Seddon, E. L. Goldsmith, “Intake Aerodynamics”, AIAA Education Series, Reston, VA, P. 274  Dr. Larry Redekopp and Dr. Dennis Phares

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x Streamline

      0.764 

5/ 4

 U 1      

1/ 4

(5)

Based on the advisor’s opinions, the following assumption was made about  , where dPitot=2.0 (6)

  Cd Pitot mm:

Substituting different values of C, and utilizing the assumptions presented in Table 1, the streamline Length corresponding to each value of C within the mentioned range of values was calculated using Equation 5. The results are presented in Table 2. Parameter Value at Sea Level  1.225 kg/m3  1.789x10-5 N·s·m-2 U1 30 m/s Table 1. Atmospheric and flow parameters

C C1=5.0 C2=7.0 C3=10.0

  Cd Pitot 0.010 m 0.014 m 0.020 m

xStreamline Rex 5 0.1676 m 3.44x10 (Turbulent) 0.2552 m 5.24x105 (Turbulent) 0.3986 m 8.19x105 (Turbulent)

Table 2. Streamline length

According to Schlichting**, the combined boundary layer travels as a free turbulent

wake after the convergence of the two ducts. Based on the explanation provided by the source, the free turbulent wake created after convergence of the two duct walls could be classified as a two-dimensional simple wake for which two main phenomena are observed: first, the increase in the thickness of the boundary layer, and second, depression of the velocity profile of the wake. It is noted in the source (Table 24.1) that in turbulent flows, an increase in the width of the boundary layer is proportional to x0.5 and a decrease in velocity is proportional to x-0.5, where x is the distance that the wake travels as a free turbulent flow.                                                             **

Schlichting, “Boundary Layer Theory”, 1968, McGraw Hill 

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Schlichting does not provide an equality for estimating the amount of depression in the velocity profile of the free wake or the increase in the width of the wake, but clarifies as long as x is relatively small compared to the characteristic length of the system, such deviations from steady flow are negligible.

This is true for the proposed model as long as the

measurements are made close to the intersection of the two ducts. Selecting a value for C between five and ten it is decided that C=7 will provide sufficient margin to measure the boundary layer (according to Equation 10 the intake exit boundary layer thickness, ď „ , is 14 mm). Based on this selection for the thickness of the boundary layer and the corresponding length of the streamline wall, the geometry of the aircraft’s intake was scaled down to a model of approximate dimensions (1x0.8x0.5) feet, with a streamline length of 0.255 m. This model was small enough to be tested in the wind tunnel available in BHE-301.

2. Experimental Techniques: 2.1 Modeling and manufacturing of the apparatus: The apparatus is designed using AutoDesk Mechanical Desktop surface editing package and is manufactured using a three-dimensional printer. The 3-view of the apparatus, as well as an

Fig. 9) Scaled model in 3-view Presented dimensions are in feet with four significant digit accuracy provided by the CAD software. Manufacturing tolerance of the 3d printer is 0.04 inches

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isometric view of the model made using the CAD software is presented in figures 9 and 10. The

apparatus

for

this

experiment

is

manufactured using a 3-d printer. Due to the limitations on the maximum printable size of the 3-d printer, the design was decomposed to 3 main pieces shown in figure 11.

Fig. 10) Rendered view of the bifurcated intake system showing the double slot pattern created for suction

Due to the brittleness of the printed parts, the team was advised to apply chemical hardeners to improve the stiffness of the model. The hardening process was repeated twice to achieve a good diffusion on the surface of the printed parts. After hardening the ducts were divided in to two pieces using the table saw to allow the sanding of the interior surfaces of the ducts to minimize the effects of the roughness of the surface on the formation of boundary layers inside the duct. As it was intended during the design, four slots

Fig. 11) Manufactured parts of the model.

of width ď „ /4 and spaced evenly in ď „ /4 intervals were integrated into the ducts (see figure 10). Extreme care was taken to ensure the structural integrity of the slots, given their relatively small size and high risk of rupture. In order to induce the pressure gradient at the intersection of the two ducts, another part was added to the assembly. As it is

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evident from figure 12, this part seals a chamber of low pressure behind the openings on the inner walls of the ducts. A Âź inch tube was selected to connect the suction pump to the apparatus. All the elements were bonded together using quick-drying epoxy and all gaps

Fig. 12) Assembly of the suction system

(out of which air could escape) were filled using additional resin. After the manufacturing of the apparatus parts was finished, a wooden base was designed and manufactured to secure the model inside the wind tunnel. In order to minimize the drag of the apparatus as well as avoiding the creation of turbulent flows, circular cross-section was selected for the supports. The minimum height of the supports was later increased to 8 inches in order to place the intake system about the center line of the wind tunnel, therefore minimizing the interference with

the

boundary

layers formed along the wind tunnel walls.

Figure 13 presents the final intake-support

assembly.

Figure 13. (a) Rendered CAD model of the bifurcated intake system on its test stand; (b) Side view of the experimental setup that was tested inside the wind tunnel.

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The intake assembly was connected to the supports using plastic zip ties to allow for a rigid and simple connection, while allowing for the possibility to remove the intake system if necessary. 2.2 The test setup and the experiments: In order to determine the velocity field at the exit cross section of the intake model, the apparatus was mounted inside the wind tunnel located at BHE-301. A movable two millimeter Pitot tube combined with a data acquisition and movement control codes written in LabView environment†† allowed the measurement of the velocity to be performed at predetermined locations to obtain the base velocity profile with the downstream air speed of 30 m/sec. For this control experiment, the suction chamber of the apparatus was isolated from the inside of the intake model via a thin tape covering the opening on the interior walls of the duct. These measurements were performed using a coordinate template (presented in

Fig. 14) the coordinate template showing the measurement locations and the predicted boundary layers.

figure 14) with 126 measurement points, of which 100 were located within the predicted dorsal boundary layer region. This decision was made considering the fact that the velocity inside the left and right side regions of the exit will most likely be constant throughout, while the dorsal regions velocity will be significantly lower than the other segments with a high velocity gradient due to the presence of the boundary layer. The pitot tube used in this experiment was installed on a movable base to control its horizontal and vertical position inside the wind tunnel test section. A Labview codes was                                                             ††

LabView Codes were provided by the laboratory as a part of the wind tunnel facilities operated at BHE.

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used to control the motion of the pitot tube inside the wind tunnel. To organize the process of measurement, a spreadsheet was prepared with the coordinated of all the data points presented in figure 14. An analog to digital converter was used to convert the analog output signal of the Honeywell pressure transducers (connected to the pitot tube) to digital signals to be transferred to the computer. By running another LabView code, the captured signals where converted to velocity values. This process was performed in a real-time manner. Before running each test, the instrumentation was calibrated in the LabView environment to report the value of the upstream and downstream air speed as zero. It was observed that in order to receive the correct velocity readings during a test run, each measurement attempt should last at least 5 seconds. After consulting with the lab manager, it was determined that the most likely value of uncertainty for the measured airspeed is ±1 m/sec. After completion of the control experiment, a series of tests were performed with the vacuum system operational to quantify the effects of the suction system on the thickness and intensity of the boundary layer. A vacuum pump available in the laboratory was used to provide the pressure difference necessary for the operation of the vacuum system. A differential pressure gauge equipped with a small gate valve was used to adjust and monitor the local pressure induced inside the vacuum hose connected to the intake model. In order to calculate the flow rate trough the vacuum system using the differential pressure gauge, Bernoulli’s equation is applied between the ambient atmosphere and the pressure gauge to obtain (7) for the flow rate through the vacuum chamber: QS 

2 2 P dtube  4 

(7)

Where P is the gage pressure and dtube is the diameter of the tube. 13


Figure 15 demonstrate a schematic of the general setup used for the experiment, with the vacuum system included.

Intake System Vacuum Chamber

Downstream Pitot tube + Pressure Sensor

P

Stagnation Pitot tube + Pressure Sensor

P Vin

GND

Vref

ENB

A/D Converter

ENB

Vout

B1

B8

Sign

GND

Vref

B8

B1

D/A Converter

Test Cross-Section of the wind tunnel

Analog to Digital Converter

Vertical Step Motor

Horizontal Step Motor

Digital to Analog Converter

Data Acquisition and Measurement control Differential Pressure Gauge Valve

Vacuum Pump

Disconnect

Figure 15) the general setup for the wind tunnel experiments. The main measurement system is in black, while the suction system is in blue. Note that the suction system was not operational for the control experiment but is added for completeness of the schematic.

Figure 16 shows the apparatus installed inside the wind tunnel as well as the pitot tube and the suction hosing prior to a test run. The model is accessed through a small circular opening as well as the top hatch of the test section, not visible in figure 16. It was determined that the vacuum pump can induce maximum of 11.85±0.01 inHg (40.02±0.03 KPa) pressure difference, therefore a decision was made to perform the vacuum Figure 16) the apparatus mounted inside the wind tunnel’s cross-section. Note the location of the experiments in 3 intermediate pressures of 2,4, and pitot tube inside the intake’s exit as well as the vacuum hose connected to the apparatus.

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8 inHg to observe the effects of the different flow rates on the thickness and intensity of the boundary layer. The obtained velocity readings from the control experiment and the 3 vacuum tests were then processed in order to provide a quantitative measure of the effectiveness of the dorsal boundary layer removal method in reducing the distortion level of the intake system. A similar procedure was followed for these experiments. Naturally, the thin tape covering the slots on the interior duct walls were removed, and the suction pressure was adjusted to the aforementioned values for different experiments. It was suggested by the project advisor that the losses induced inside the gate valve and the hosing might lead to error in flow rate calculations, and therefore alternative methods to measure the flow through the suction chamber was explored. The only available alternative, using a wet test cell flow meter was proven impossible, since both machines available on campus were only capable of performing measurements on an outgoing flow and incompatible with suction systems. In order to generate a visual representation of the collected data, MAPLE data processing module was used to generate vector plots of the airspeed based on the values obtained during the wind tunnel experiments. Also, the curve fitting package of the same software was used in order to fit a 5rd order spline to the data obtained along the horizontal centerline of the inlet exit. The disturbance thickness of the boundary layer at the horizontal center line was measured using the fitted curves to the data obtained experimentally. The spreadsheets containing the data obtained from the experiments were also processed in order to compute the average distortion coefficient for each test case.

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3. Results and Discussion: Figures 17 trough 21 presents the results of the experiments performed in form of vector plots of the airspeed in a cross-section 10 millimeters away from the duct intersection plane. Figure 17 present the results for the control experiment, while figures 18-21 demonstrate the effects of the suction flow (QS) on the boundary layer’s velocity field.

Fig. 18) Velocity vector plot, QS=3.33 cm3/s.

Fig. 20) Velocity vector plot, QS=6.65cm3/s.

Fig. 17) Velocity vector plot, QS=0 cm3/s.

Fig. 19) Velocity vector plot, QS=4.71cm3/s.

Fig. 21) Velocity vector plot, QS=8.10cm3/s.

It should be noted that the coloring of the vector plots is merely used to identify the lateral position of the data points. The blue, red, orange and magenta colored vectors are located 16 Â


respectively 3±0.2, 6±0.2, 12.3±0.2, and 20.3±0.2 millimeters away from the center respectively. As it is evident from figure 17, there is a noticeable change in the velocity magnitudes in the dorsal region of the intake exit in the absence of the suction. It is also noticed that throughout the dorsal region, the velocity increases slightly towards the bottom of the intake. This could be resulted from inaccuracies in the manufacturing of the apparatus or unknown turbulent behavior initiated by the models curve. Comparing figures 17 to 21, it is obvious that the average velocity of the dorsal region increases as a higher suction flow rate is applied to the suction chamber installed at the intersection of the two ducts. In order to better observe the effects of the application of suction on the velocity profile of the boundary layer, a 4th order spline was fitted to the data along the horizontal diameter of the cross section. These velocity profiles are presented in figures 22. An enlarged view of the graph limited to the dorsal boundary layer region is also presented in figure 23.

Fig. 22) velocity profiles along the horizontal centerline for different suction flow rates.

Fig. 23) an enlarged view of the dorsal region of the velocity profiles presented in Fig. 23

From figure 23 it can be seen that no significant change in the boundary layer velocity profile is observed as a result of a suction flow rate of 3.3±0.01 cm3/s, while a significant increase in

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the average velocity of the boundary layer is observed associated with the suction flow rate of 4.7±0.01 cm3/s. For all design purposes, it seems beneficial to compute the ratio between the suction flow rate and the total flow rate through the intake system to have an idea about the feasibility of the proposed solution in a real world design scenario. Also, in order to quantify the influence of the suction system on the behavior of the boundary layer, it is possible to define the non-dimensional free wake ratio (  ) as the following:



U U

(8)

Where:U: Average velocity on both sides of the boundary layer. and ΔU: Depression observed in the boundary layer region.

Table 4 presents a summary of the results of the analysis performed using the experimental data for the velocity profiles and the corresponding suction flow

Fig. 24) Nomenclature for the free wake ratio

rates (QS). For these calculations the total flow through the intake (Q0) is calculated based on the average air velocity at the exit and the exit area of the model. QS Q0 (%)

 (mm)

 (%)

0

QS [ cm3 s .] 0

0

11±2

20.2

6.75±0.01

3.33±0.01

0.004

10±2

17.7

13.51±0.0 1 27.01±0.0 3 40.02±0.0 3

4.71±0.01

0.005

9±2

14.3

6.65±0.01

0.007

10±2

9.4

8.10±0.01

0.009

12±2

5.1

P S [KPa]

Table 4. Suction flow rate (QS), flow rate ratio (QS/Q0), displacement thickness (Δ), and the free wake ratio (  ) corresponding to different gauge pressures.

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Reviewing the results presented in this table, it is concluded that by inducing a flow rate approximately 0.01 percent of the total air flow through the intake, the proposed vacuum system can reduce the free wake ratio by 71 percent, therefore exceeding the initial goal for the project (70 %). Such a large reducing in the non-dimensional wake ratio could reduce the magnitude of the cyclic forces imposed on the fan blades tremendously. Another popular approach to assess the distortion properties of the flow through the intake system is to calculate the distortion coefficient (DC) for the intake. This parameter also will provide a measure of the non-homogeneity of the velocity profile. Based on the theory presented by Seddon et al.‡‡, the distortion coefficient for an arbitrary sector of the intake exit is a measure of the stability of the flow in the intake and is defined as the difference between the maximum and minimum velocity divided by the average velocity in the sector. Equation 9 shows the relation for the distortion coefficient: V Vmax - Vmin  V V

Eqn. (9)

For n number of computational sectors, the average distortion coefficient is calculated from:

V V DC  i 1 n n

Eqn. (10)

  The computational sectors used to calculate the distortion

coefficients are shown in figure 25. This template is consisted of 36 unequal panels. The calculation has been done using an Excel spreadsheet containing the values of the velocity vectors separated into the panels shown in figure 25.

Fig. 25) Computational sectors used for calculating the distortion coefficients

                                                            ‡‡

J. Seddon E. L. Goldsmith, “Intake Aerodynamics”, AIAA Education Series, Reston, VA, P. 274 

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The distortion coefficients were calculated for all test cases and the values are presented in table 5. QS [cm3 s .]

DC

0 3.33±0.01 4.71±0.01 6.65±0.01 8.10±0.01 0.098 0.094 0.091 0.105 0.102

Table 5. Distortion Coefficients corresponding to different values of suction flow rate.

It can be seen that the distortion coefficient is reduced by approximately 13 percent for a flow rate of 8.1±0.01 cm3/sec. Based on one of the faculty advisors opinion§§, given the extremely importance of the distortion levels for modern small size engines, a 13 percent reduction in the distortion created by an intake system could be very significant, ultimately allowing a larger number of jet engines to be considered for bifurcated intake integration.

4. Conclusion As it was shown in this report, the internal boundary layer removal is believed to be a feasible solution to resolve the issue of the dorsal boundary layer created in bifurcated intakes. It can be argued that the reduction of the boundary layers along the inner walls of a bifurcated intake system could be both financially and operationally beneficial. In this report, it was shown that the system is able to improve the non-dimensional wake ratio by approximately 70 percent with a suction flow rate of only 0.01 percent of that flow passing through the intake itself. Although one may argue the issue of weight as a disadvantage of the suction device, as well as an increase in price of the development of the engine and design of the intake system, the issue is compensated by the improved reliability and durability of the engine and, as a result, the aircraft. The need for engine overhauls, maintenance costs, and the                                                             §§

Mr. mark Page, USC faculty member and senior scientist at Swift Engineering Co.  

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risk of serious accident could decrease as a result of application of this system. The suction needed could be provided by means of using a standard vacuum pump installed on the power plant*** which reduces the design cost and complexity of the final product. Such a system is presented in figure 26. Since a 70 percent reduction in the wake ratio was achieved it can be argued that the initial goal for the project was reached.

Given the complexity of the design and

manufacturing phase as well as the time necessary to complete all the testing and data processing, it is notable that the project progressed as scheduled with very minor delays. The most important improvements necessary in the project is utilization of a more accurate method for measuring the flow through the suction chamber to verify the values obtained for the suction flow rates by applying Bernoulli’s principle. Tests with higher air speeds or at non-zero attack angles may provide insight towards a more optimized solution for the internal

boundary

layer

removal

concept,

particularly with regard to the shape and the arrangement of the suction slots. Application of porous materials instead of open slots is also of interest due to their capability to cause gradual reduction in the thickness of the boundary layer.

Fig. 26) A top view of the suction chamber device and co-axial vacuum pump in realworld powerplant integration.

                                                            ***

This unit would use the mechanical power provided by the gear box of an installed power plant.

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References - R. W. Fox, P.J. Pritchard, A.T. McDonald, “Introduction to Fluid Mechanics 7TH Edition”, 2009, John Riley & Sons - J. Seddon, E. L. Goldsmith, “Intake Aerodynamics”, AIAA Education Series, Reston, VA, - Schlichting, “Boundary Layer Theory”, 1968, McGraw Hill

Appendix A: A-1 Real Size Intake Design Problem: AC

Inlet Frontal Area

 Turb KBPR  

Boundary Layer Thickness Bypass Ratio Factor Critical Diameter (Diameter of Intake Exit) Total Air Mass Flow Rate at Engine Intake

L

m  a        m  cool   m  gas  

Air Mass Flow Rate Required for Engine Cooling Air Mass Flow Rate Required for Engine Combustion

N eng

Number of Engines Air Density

 TTO U1 xStreamline

Takeoff Thrust Steady State Speed Streamline Length Adjacent to Duct Wall

In order to design the experimental setup, the request for proposal (RFP) of the 20072008 AIAA Individual Aircraft Design Competition was reviewed. This RFP asks for a realworld aircraft. The aircraft in response to the RFP uses a bifurcated intake system that was designed based on the method presented by Jan Roskam†††. The intake system for this aircraft has been optimized for operating at the cruise condition presented by the RFP (V = 350 kts, Alt=35000 ft). Based on Roskam, the required frontal inlet area for a subsonic jet engine is calculated from:                                                             †††

Roskam J., Airplane Design Part VI ; Section 6.2 PP 147-182; 1990

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AC =

a m U1 ρ

(A-1)

Where the total air mass flow rate at the engine intake is computed from:

m  a =m  gas +m  cool

(A-2)

The air mass flow rate per engine required for engine cooling is found from:

m  cool = 0.06 m  gas

(A-3)

The air mass flow rate per engine required for the proper combustion is found from:

m  gas = KBPR

TTO N eng

(A-4)

Based on the result of this calculation, the geometry for the air intake is developed in order to achieve a reasonable rate of change in the area of the intake cross-section. The surface model for the bifurcated intake was prepared using the surface editing package

U1 M1 ρ35,000

350 kts. 0.587

 gas m

1.08 slug s .

m  cool m a AC

0.06 slug s .

0.00062 slug ft 3

1.14 slug s . 2.91 ft 2

within Autodesk Mechanical Desktop. The result of the lofting and Table normal on A.1 adjustment Intake Parameters of full-size aircraft

the CAD model can be seen in figure 4 below.

Fig. A-1) the real-size bifurcated intake designed as a response to the AIAA

23


2007-2008 AIAA Individual Aircraft Design Competition Request for Proposal

A- DESIGN REQUIREMENTS AND CONSTRAINTS: 3.1 Design Mission 3.1.1 General Design Requirements

The VLJ Trainer will be capable of carrying two people side-by-side. No ejection system is required or allowed. The aircraft should be able to land within 2,000 ft field length at maximum take-off weight The aircraft must have a range of at least 800 nm (full fuel, maximum payload). The aircraft must possess a top speed of at least 350 kts and have a service ceiling of at least 36,000 ft. Clean stall speed may not exceed 70 kts at sea-level at maximum take-off weight. Landing gear must be retractable. Payload:

Two 220 lb pilots and 50 lb baggage. 3.1.2 Required Mission Performance

1. Warm-up and taxi at idle power for 8 minutes. 2. Take-off fuel allowance is equal to the fuel consumed during 2 minutes of operation at maximum take-off power. 3. Take-off field length must not exceed 2,000 ft. on a hot day (95oF) at sea level. 4. Cruise/climb to cruise altitude of 35,000ft at 3,000 fpm. 5. Cruise at 35,000 altitude and 350 kts. 6. Descend to 1000 ft for 100 nm. 7. Loiter for 45 minutes (reserve) 7 Descend to sea-level. 8 Land (distance not to exceed 2,000 ft at max. take-off weight) and taxi.

24 Â

Reduction of Distortion in Bifurcated Intake Systems Utilizing Boundary Layer Ingestion