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Theoretical and Experimental Modeling of Vortex Engine in Ramjet Application

S. Besharat Shafiei1 and A. Ghafourian2

Sharif University of Technology, Tehran, Iran, 11365-8639

M. H. Saidi3 Sharif University of Technology, Tehran, Tehran, 11365-8639, Iran

and

A. A. Mozafari4 Sharif University of Technology, Tehran, Tehran, 11365-8639, Iran

A new experimental facility was designed, fabricated and tested to model and study the possibility of applying the bidirectional swirl flow on the combustion chamber of airbreathing subsonic ramjet engine. Appropriate intake was designed to convert axial external air to tangential swirl flow inside the combustion chamber. Inlets with appropriate angles conduct the swirl flow into the chamber and create bidirectional swirl flow field in the combustion chamber. This flow field has been modeled theoretically to determine the velocity field characteristics by previous researchers. Bidirectional swirl flow in liquid fuel ramjet engines has the proven advantage of keeping the combustion chamber walls cool and in solid fuel ramjet engines increases the fuel burning rate as demonstrated in the previous investigations. The experimental study was performed by using propane and air as fuel and oxidizer, respectively. The temperature of combustion chamber wall was measured and found to be as low as 100 . The possibility of combining the airbreathing bidirectional swirl flow and ramjet engine was proved experimentally. The temperature variations at 3 points of the combustion chamber wall of this airbreathing bidirectional swirl flow engine were measured relative to time. The temperature variations on the same three points relative to various mass fuel air ratios were measured and are reported. Further development of this type of combustion chamber enables manufacturers to use less expensive and more available material in their production of combustors with improved combustion efficiency due to the increased mixing rate and cooled combustion chamber wall.

Nomenclature l : Chamber aspect ratio L :Chamber length

FuelD :Inlet fuelpipediameter .

Airm :Air mass flow rate .

fuelm :Fuel mass flow rate

P :Chamber pressure

0Q :Outlet volumetric flow rate

iQ :Inlet volumetric flow rate R :radialcoordinate

1 Graduate Student, Aerospace Engineering, besharat@alum.sharif.edu, AIAA Student Membership. 2 Professor, Aerospace Engineering, Ghafourian@sharif.edu, No membership. 3 Professor, Mechanical Engineering, Saman@sharif.edu, No Membership. 4 Professor, Mechanical Engineering, Mozafari@sharif.edu, No Membership.

45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit2 - 5 August 2009, Denver, Colorado

AIAA 2009-5433

Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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wallT :Chamber wall temperature

rU :Radial velocity

zU :Axial Velocity u :Tangential velocityθ U :Inlet velocity Z :AxialCoordinate

:Densityρ :Equivalence Ratioφ

b / aβ =

I. Introduction Manipulation of flow field to reduce the heating rate of the walls has been used for many years. The idea using excessive oxidizer or keeping the combustion wall cool has been used by combustion chamber designers. These methods may result in reduction of combustor performance or increase in the overall weight of the system. Extensive efforts have been made by researchers to investigate the swirl flow field and its advantages and important parameters. Chiaverini et. al. [1] worked on vortex thrust chambers employing an oxidizer swirl injector just upstream of the converging section of nozzle to generate a coaxial vortex flow field in their combustion chamber. They found substantial reduction in chamber wall temperature. Vyas, et. al. [2] derived an exact solution for the flow field of a bidirectional coaxial vortex for a liquid propellant combustion chamber. This study proved that the flow does not short circuit from the injection point to the exit nozzle, but rather a tendency exists for the flow to go all the way up one end and returns back through a core flow, which is referred to as a bidirectional flow. Jahangirian, et.al. [3] studied the effect of bidirectional swirl flow on the rate of heat transfer to combustion walls. They experimentally concluded that in cases with bidirectional swirl flow, highest efficiency and lowest wall temperature there exist. This can be due to better mixing of fuel and oxidizer and absence of hot spots in the combustion core. Akbari, et. al. [4] investigated flow field in an axisymmetric laboratory combustor. Their results indicated that the instability frequency was controlled by the inlet pipe acoustics and the vortex convection inside the combustor. In this study the concept of using bidirectional swirl flow in combustion chamber of a subsonic airbreathing ramjet engine was proved experimentally. The best geometry for an intake to capture axially flowing external air relative to combustion chamber and converting it to bidirectional swirl flow inside the combustion chamber has been designed. The appropriate angle for installing the fuel injectors to minimize the wall temperature was determined numerically. Two disadvantages of the subsonic ramjet engines are higher wall temperature and instability, in which by using swirl flow in combustion chamber, these two problems can be solved as will be explained in this research.

II. Theoretical Background

Several researchers in the swirl flow field have analyzed the combustion field with swirl flow and have solved the conservation equations with respect to boundary conditions and have determined many parameters of this field. Figure 1 shows a typical geometry for a chamber with a closed head and open end analyzed by Bloor and Ingham [5].

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Figure 1. Combustion chamber geometric parameters The appropriate intake should be located at the end in order to convert axial air flow to tangential on the inner layer of chamber. Intake cross sectional area should be such as the air does not enter the combustion chamber with radial or axial components, but just tangentially. By ignoring the very small transition zone at the intake to the chamber, the rest of the flow is axisymmetric. The analysis can be done for both compressible and incompressible flows, but in this case the incompressible flow was analyzed. By assuming: steady, inviscid, incompressible, rotational, and non- reactive, the governing equations can be simplified as [2] :

r z(ru ) u1 0r r z∂ ∂

+ =∂ ∂

(1)

2r r

r zuu u 1 pu u

r z r rθ∂ ∂ ∂

+ − = −ρ∂ ∂ ∂

(2)

rr

u u uu 0

r rθ θ∂+ =

∂

(3)

z zr z

u u 1 pu ur z z

∂ ∂ ∂+ = −

ρ∂ ∂ ∂

(4)

Applying the following boundary conditions these equations can be solved:

r a, z L, u Uθ= = =

(5)

zz 0, r, u 0= ∀ =

(6)

rr 0, z,u 0= ∀ =

(7)

rr a, z,u 0= ∀ =

(8)

0 i iz L,Q Q UA= = = (9)

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Solutions to the above equations and the axial velocity (uz) in the combustion chamber show the existence of a bidirectional swirl flow.

III. Experimental Procedure An axisymmetric combustion chamber with a suitable intake has been designed and fabricated. In the way of reaching the best geometry for intake to be able in producing a swirl flow field, we fabricated some geometries and performed hot tests of them. From the analysis of the performance of different geometries, consequently best geometry for intake and swirl airbreathing engine with this flow field was fabricated and experimentally investigated to measure its variable. Figure 2 shows a preliminary geometry for intake that is able to produce swirl flow field however due to high pressure loss is not appropriate to use in an airbreathing engine. Figure 3 indicates the profile of the selected combustion chamber and intake. Combustion chamber specifications there are in Table 1. The proposed intake is able to convert directional air flow to tangential one. The aforementioned profile does not have high pressure loss and drag compare to other designs. Furthermore, this intake can be folded and this advantage is a relevant property in aerospace industries. Figure 4 shows a view of the board with some subsystems used.

Figure 2. Shape of preliminary profile of the model

Figure 3. Profile of final tested model

Table 1. Combustion chamber specification

Dimension Value Parameter m 0.075 a m 0.3 L - 0.7 β

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Figure 4. View of board used

Combustion chamber body includes one intake at chamber end. At ideal performance it would be better to have 2 intakes at chamber end to have a symmetric geometry. Two injectors are used at the chamber head with an angle of 15۫ relative to the chamber axis. The above mentioned angle is selected from examining different angles, such as 60, 30, 15, 0 degree for injector relative to the chamber axis and found that in 15 degree, wall temperature is lower than that of other angles. Geometric parameters of the chamber are chamber length (L), exhaust nozzle diameter (b) and injector diameter (Dfuel). The chamber is made of “ST37” material. The chamber has an inside diameter of 15.4 cm. Based on Eric Rice [6] research variation of chamber diameter does not have much influence on the result. The length of chamber is 28 cm (L=28cm). The intake cross sectional area is 28.26 cm2 that is equal to the fan exit area. Also an opening with the area of 28.26 cm2 should be cutted on the wall chamber till the captured air by intake enter the chamber, this opening should be designed carefully to be able to convert the axial flow to tangential one successfully. We used a rectangular opening with length of 10 cm and width of 4 cm for this cutted part. This arrangement tends to cause have swirl flow field in the combustion chamber and stable combustion occurs in the chamber. In the case of triangle or circle or oval geometry of opening, tests showed that bidirectional swirl and stable flame does not exist or in some cases combustion occurred out of combustion chamber. The serious maloperation of the engine is due to the shape of opening which is not appropriate such that be able to convert captured direct air to tangential flow and in some cases captured air goes out of combustion chamber by nozzle. Having selected rectangular opening, the entrance air has tangential component just after entering. Ignition starts with an external means. Oxidizer feeding system comprises of a fan that supplies the required air as oxidizer air to the intake. In the case of combining the axial and swirl flows, a compressor with a maximum absolute pressure of 7 bar provides the air supply for intake and fan provides the air for axial entrance. In this case a combination of filter-regulator controls pressure of the air line. Fuel feeding system supplies the fuel to the chamber using 2 injectors, each of them with 0.6 cm diameter. Propane is used as fuel. A calibrated rotameter is used to measure fuel flow rate. Propane is injected into the chamber with 16 m/sec and the profile of fan speed is presented in Fig 5. The average

fan speed is approximately 8 m/sec. .

fm 0.0017 kg / sec= and 2fA 0.0556cm= , Therefore we used two injectors

with 0.6 cm diameter. A “K” type thermometer is in the chamber wall to measure wall temperature.

IV. Experimantal Sequences

Figure 5. Velocity Profile of fan

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V. Experimental Sequences The first type of experiments were performed to verify the operating limits of combustion chamber, measuring and controlling the parameters. These parameters are in Table 2. Operating limits are in Table 3. Then swirl flow airbreathing subsonic ramjet chamber was tested and the variation of wall temperature v. s. time in three points namely the head, middle, and top of the combustion chamber was measured. Results are in Table 4.

Table 2. Controlled and measured parameters

To be measured

To be controlled

Row

Tch &Pch .fuelm

1

Tch &Pch .Airm

2

Table 3. Mass Flow rate limits

Table 4. Wall Temperature & Combustion chamber static pressure in various fuel /air ratios

Max. amount

Min. Amount

Dimension Inlet fuel rate

Engine Type

11.2 0.3 gr/sec .fuelm

Axial

38 0.3 gr/sec .Airm

Axial

8.33 0.05 gr/sec .fuelm

Bidirectional vortex

26 0.3 gr/sec .Airm

Bidirectional vortex

9 0.1 gr/sec .fuelm

Axial+Bidirectional Vortex

42 0.3 gr/sec .Airm

Axial+Bidirectional Vortex

fm.

Pfan

(mm-H2O)

Pcombustion

(mm-H2O) 1T

2T

3T

af mm..

/

5

(g/sec)

29 160 50

©

52

©

55

©

1.25

6.67

(g/sec)

33 190 51 58 58 0.47

7.634

(g/sec)

39 250 60 62 68 0.382

8.33

(g/sec)

50 310 68 75 80 0.32

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VI. Results and Discussion The chamber wall temperature variation of vortex engine v. s. mass fuel/air ratios is presented in Fig. 6. These temperatures were measured for the profile shown in Fig. 3, while the flame in the combustion chamber was presented. Air entering from the engine end with a tangential direction relative to the chamber wall moves in a swirl flow toward the chamber head and returns toward the nozzle end keeping its swirl flow direction and establishes a core flow inside the outer swirl flow. At any cross section of the chamber the direction of flow on the circumference is opposite of the flow direction in the center. This can be seen from the pyrex glass window installed in the chamber wall.

Figure 6. Temperature v. s. Fuel-air ratio in 3 points on the combustion chamber wall of vortex engine. T1: 5

cm from head, T2: 12 cm from head, T3: 20 cm from head.

As Fig. 7 and Fig. 8 indicate the chamber wall temperature is much lower in the case of combination of bidirectional swirl and axial flows as compared to the case where both fuel and air enter axially from the chamber head. For the given flow rate and chamber geometry used for these tests, the chamber wall temperature was lowered approximately 75% when bidirectional swirl flow was used. Test of the engine compound of tangential and axial oxidizer proved that the combination of two flow patterns causes a wall temperature reduction of combustion chamber much lower than the vortex case with 80% compared with the axial case. This advantage is because combustion was most complete in this case as compared to the just tangential oxidizer case.

Figure 7. Combustion chamber wall Temperature v. s. time in the middle length of compound (vortex and axial oxidizer flow) engine, in Ф~1.

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Figure 8. combustion chamber wall temperature v. s. time in the middle of axial flow engine Ф~1. It is reasonable to assume that heat transfer from the combustion core to the chamber walls is done by radiation. Cooling of chamber walls are done by convection. Therefore, in the compound case, a better chance of mixing there exists as compared to the just tangential case. This resulted in a distributed combustion through the chamber length.

VII. Conclusion Based on analytical results for the existence of bidirectional swirl flow in certain chamber geometries with controlled oxidizer and flow condition and with the motivation of keeping combustion chamber walls cool, a new experimental facility was designed, fabricated and tested. In this effort an appropriate intake was designed to be able to capture free air with least pressure loss, converting axial air to tangential one. This intake was used to provide oxidizer for a subsonic airbreathing ramjet engine. Propane and air were used as fuel and oxidizer, respectively. In case of existing bidirectional flow, wall temperature reductions of up to 75% were observed. In experiments where only some of the oxidizer was injected from the chamber end, to generate the bidirectional swirl flow and some of the oxidizer was injected with the fuel from the chamber head, the best combustion efficiency and lowest wall temperature was observed. Further development of this technique may result in reduction of instabilities in combustion of ramjet engines and cost saving for constructing this engine.

Acknowledgments S. Besharat Shafiei, A. Ghafourian, M. H. Saidi, A. A. Mozafari thank for financial support of Sharif University of Technology.

References

1 Chiaverini, M. J., Malecki, M. J., Knuth, W. K., Gramer, D.J., ” Vortex Thrust Chamber Testing and Analysis for O2-H2 propulsion Application,” AIAA paper 2003- 4473, July 2003.

2 Vyas, A. B., Majdalani, J., and Chiaverini, M. J., ” The Bidirectional Vortex. Part 1: An Exact Inviscid Solution,” AIAA paper 2003-5052, July 2003.

3 Jahangirian, S., Abarham, M., Ghafourian, A., Saidi, M. H., “Effect of Vortex Flow on Heat Transfer to Combustion Chamber Wall,” ASME Conference and Exhibit, California, USA, 2004.

4 Akbari, P., Ghafourian, A., Mazaheri, K., ” Experimental Investigation of combustion Instability in an Axisymmetric Laboratory Ramjet,” 35th AIAA/ASME/SAE/ASEE Joint propulsion conference and Exhibit, Los Angeles, USA, 1999, pp. 1-10.

5 Bloor, M. I. G., and Ingham, D. B., ” The flow in industrial cyclones,” Fluid Mechanics Journal, Vol. 178, 1987, pp. 507-519. 6 Amanpour, H., “Instability Analysis in Liguid Fuel Vortex Engine,” Sharif University of Technology., M. Sc. Thesis, 2008.