Performance improved by multistage rockets ejection in RBCC engine

ZHANG Man1, HE Guo-qiang2 and Liu Pei-jin3

National Key Laboratory of Combustion, Flow and Thermo-Structure, Northwestern Polytechnical University, 710072 Xian，People’s Republic of China

Three-dimensional two-phase numerical method is used to investigate the combustion processes in RBCC ejector mode. A new structure of primary rockets arrangement method namely multistage rockets ejection is proposed, that is to put one primary rocket in combustion chamber embedded in the rearward-facing step located at the 50% distance of upper longitudinal wall in combustion chamber, and left the other primary rocket embedded in the center strut. The performance of RBCC investigated shows that multistage rockets ejection has improved performance compared to ordinary rockets structure. At the state of 1st stage rocket working at equivalent ratio and the 2nd stage rocket of fuel-rich, the thrust of the whole engine could be increased associated with ejection mass flow rate also increased. The results with deeply analysis show that the 1st stage rocket not only ejects air coming from far-field, but also suppresses the chamber pressure affecting the inlet. The 2nd stage rocket ejects the flow in chamber continuously, and works as the flame holder simultaneity. The mechanism of multistage rockets ejection is the optimized combination of DAB and SMC. It may be the better choice under current research of the RBCC engine in ejector mode for secondary combustion.

I. Introduction BCC (Rocket Based Combined Cycle) is an advanced propulsion system concept proposed in the 1960’s. The concept combines advantages of conventional rocket and air-breathing motor to form a kind of multi working

mode (including rocket ejector, ramjet, scramjet and pure rocket)1. It is an ideal propulsion engine of reusable launch vehicle, which could greatly reduce cost and improve efficiency of propulsion system. Primary rocket is one of the basic components of RBCC, and is the key point making RBCC different from other propulsion system. In different mode, the primary rocket works in different manner. In rocket ejector mode, especially in low flight velocity phase, primary rocket is the main factor to ensure motor thrust; with flying velocity increased, primary rocket could be used as flame pilot/holder equipment, or stop working.

R

The secondary combustion is the usable way to improve the performance of RBCC engine. There are three secondary combustion organized method in RBCC ejector mode: SMC(Simultaneous Mixing and Combustion), DAB(Diffusion and Afterburning) and SPI(Shielded Primary Injection)2-4. For the first two modes, DAB mode is the most efficient combustion method, but in experiment, the exactly fuel injection direction and the secondary combustion organization is hard to carry out. In SMC mode, the incoming air mixed with fuel rich rocket plume could easily ignited, but the combustion chamber pressure should immediately affect the inlet of the whole engine.

In order to optimize the RBCC performance, a new rocket arrangement structure is discussed. Based on the former investigations in National key Lab. of Northwestern Polytechnical University, this paper carried out numerical computations to investigate multistage rocket ejection in RBCC ejector mode, that is to put a primary rocket in combustion chamber, and left one primary rocket embedded in the center strut, and analyze the performance with relative secondary combustion organized method.

1 Ph.D. student, College of Astronautics, zm2001er@yahoo.com.cn. 2 Director of National Key Lab., College of Astronautics. 3 Professor, College of Astronautics.

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II. Computational Methodology

A. Basic Solution Combustion process in rocket ejector/ramjet mode included fuel injection, atomization, vaporization, mixing and

combustion. Spray combustion flow was a complex process including mass transfer, heat transfer, chemical reaction, turbulence and two-phase flow, and related to temporal and spatial coupling. Mainly included: gas phase flow described by continuum medium conservation law, interaction effects of different ingredients, combustion, discrete phase movement and phase change, mass and heat transfer between discrete phase and continuum phase.

In this paper, solution of RANS is used SST two-equation turbulence model. 2nd-order precision upwind scheme is selected for advection term. All the numerical time steps are unified to 10-7 s. The atomization and evaporation process of the liquid fuel determine the combustion response time. For the medium discrete phase numerical terms in two-phase flow, it is widely used Euler-Lagrange particle tracking method. For spray combustion in high speed chamber, liquid fuel suffers from primary atomization process through nozzle and aerodynamic shear atomization(secondary atomization). The initial size distribution of the droplet at the point of injection nozzle has been directly taken to be 10 mμ following Rosin-Rammer distribution with the exponential factor of 3. After the liquid fuel injected into mainstream, droplets suffers from intense shear action. At the condition of critical Weber number far greater than 102 magnitudes, Reitz secondary atomization model is more appropriate.

Kerosene fuel is the mixture, commonly used by a single component of alternative fuel C12H23 in simulation. The thermal properties of each component are given in terms of polynomial fits of temperature. All coefficients (a1k~a7k) of the polynomial fits needed are taken from JANNAF data.

B. Meshes and Boundary Conditions Numerical chamber modeling is based on three-dimensional entities. Since CFD method is been remarkable

influenced by the artificial grid division, high-precision grids should be carried on the domain. The total grid cells number is 400,000 with the average wall y+ of 50. Figure 1 is the meshes of the computational domain. Numerical computing parallel platform is composed of five Pentium 3.0GHz CPU.

Fig.1 The meshes of the computational domain

The aerodynamics of incoming air is reflected on the supersonic inlet boundary. The incoming velocity, direction,

temperature and pressure are set on the far-field of the vehicle. Total pressure and total temperature are set on the rocket nozzle to simulate rocket chamber pressure. The no-slip and adiabatic boundary condition is employed on the wall.

C. Numerical Method Validation In order to verify the validity of numerical methods, this paper carries on comparison between the direct-connect

ground experimental results with relative CFD simulation. Details geometric parameters and combustion condition see the literature5. Figure 2 gives experimental pressure distribution on the chamber wall compared with numerical results.

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x /m

P/M

Pa

-0.6 -0.2 0.2 0.6 1 1.40.4

0.6

0.8

1

1.2

Sidewall,CFD This workSidewall, Exp.Upwall,CFD This workUpwall, Exp.

4.5%

[5]

[5]

Fig.2 Comparison of experimental pressure distribution with numerical prediction

Chamber pressure distribution can reflect the heat released location with combustion, in Fig. 2, good agreement

with experimental data, indicates that the numerical method is reliable, and can be more accurate description of the turbulent combustion details.

III. The Concept of Multistage Rockets Ejection The research configuration of RBCC engine is 2-D strut structure, including the fore body of vehicle, air inlet,

isolator section as well as the expansion geometry combustion chamber. The combustor has an entrance cross section of 70x100mm. It is composed of two expansion sections. The first expansion section is of 800mm length and had 2-deg divergence, whereas the second expansion section was 200m long with 4-geg divergence. Kerosene for secondary combustion is injected normal to the midstream via four orifices of 0.8mm diameter at the 65% along the flow direction of chamber. Two primary rockets are embedded in the central strut symmetrically from top to bottom. Combustion chamber geometry and primary rockets structure is shown in Fig. 3.

Isolator

Vehicle Forebody

Combustion chamber

Rocket

Block

Secondary Fuel Injection

Isolator

Vehicle Forebody

Combustion chamber

2nd Stage Rocket

Block1st Stage Rocket

Fig.3 The sketch of ordinary rockets Fig.4The sketch of multistage rockets ejection

On the basis of the ordinary structure of primary rockets mentioned above, a new structure of primary rockets is

proposed here, that is to put one primary rocket in combustion chamber embedded in the rearward-facing step located at the 50% distance of upper longitudinal wall, and left the other primary rocket embedded in the center strut. The rocket in chamber is set as the 2nd stage rocket, and the rocket left in the strut set as the 1st stage rocket. Both of two rockets constitute the multistage rockets structure, as shown in Fig. 4. In order to contrast the performance of

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multistage rockets with the ordinary primary rockets structure, keep other geometric parameters of RBCC engine as the same. For ease of the indication, supposes ordinary primary rockets structure is CASE1, and multistage rockers structure is CASE2.

IV. Results and Discussion

A. 1st stage rocket of equivalent status, 2nd stage rocket of fuel-rich The secondary combustion organized under ejector mode of RBCC, is the only feasible way to obtain the largest

thrust. In the structure of CASE1, DAB mode with kerosene spray and the SMC mode additional with rich fuel in the rockets plume for secondary combustion are carried on respectively. The total amount of jet fuel in DAB mode is equal to that of rich fuel in rockets of SMC mode. In SMC mode, the equivalent ratio is set to 1.5. In order to contrast ,in multistage rockets ejection structure of CASE2, the 1st stage rocket works at equivalent status, and the 2nd stage rocket is fuel-rich for equivalence ratio of 1.5.

Figure 5 is the RBCC engine thrust changing with the secondary fuel mass flow rate. In rocket fuel-rich condition, secondary fuel mass flow rate refers to the amount of rich fuel in rocket plume. Its increase is realized through raising rocket chamber pressure.

Fuel mass flow rate/ (Kg/s)

Thru

st/N

0 0.05 0.1 0.15 0.2 0.250

300

600

900

1200

1500

1800

CASE1, DABCASE2 ,CASE1 ,SMC

22%

Ma=0.8 , H=0km

Fuel mass flow rate/ (Kg/s)

Thru

st/N

0 0.05 0.1 0.15 0.2 0.25

600

900

1200

1500

1800

2100

2400

CASE1, DABCASE2 ,CASE1 ,SMC

19%

Ma=1.8 , H=10km

Fig.5 With secondary combustion, the relationship between fuel mass flow rate and thrust

It can be seen from Fig. 5, the thrust force of CAS2 is situated between the DAB and SMC of CASE1. Take the

flight Mach number of 0.8, secondary fuel mass rate of 0.1kg/s condition as the example, the thrust force of CASE2 is lower than the DAB condition of CASE1 22%, but upper than that of SMC mode. When the flight Mach number is 1.8, the thrust force of CASE 2 tends to be consistent with the DAB mode of CASE 1 along with fuel mass flow rate increase.

O2.Mass_Fraction: 0.02 0.06 0.1 0.14 0.18 0.22

Fig.6 Distribution of O2 in combustion chamber

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For an in-depth analysis of the flow detail and combustion efficiency, Fig. 6 is O2 mass fraction distribution in chamber of CASE 2 under the condition of Mach number 0.8 as well as primary rocket chamber pressure of 4.0 Mpa.

From Fig. 6 can see, as the 2nd stage rocket embedded at the upside wall of the combustion chamber, rich fuel in the rocket plume can not react with the ejected air completely. The flame is established near the upside wall, so that the combustion efficiency is not high. This causes the overall thrust cannot reach the maximum.

B. 1st stage rocket of equivalent status, 2nd stage rocket of fuel-rich with kerosene spray combustion The secondary combustion is modeled under the conditions that 1st stage rocket of equivalent status, and the 2nd

stage rocket of equivalence ratio 1.5, with additional kerosene spray combustion. Figure 7 is the RBCC engine thrust changing with the secondary fuel mass flow rate. Figure 8 is ejection mass

flow rate changing with the secondary fuel mass flow rate. In order to show that thermodynamics of combustion in chamber clearly, Fig. 9 is the mass fraction of O2 distribution in chamber.

2nd fuel mass flow rate/ (Kg/s)

Thru

st/N

0 0.05 0.1 0.15 0.20

300

600

900

1200

1500

CASE1, DABCASE2, spary combustionCASE2, rocket rich fuel

Rocket pressure =4.0Mpa

Rocket pressure =1.5Mpa

Ma=0.8 , H=0km

7%

15%

2nd fuel mass flow rate/ (Kg/s)

Thru

st/N

0 0.05 0.1 0.15 0.2400

800

1200

1600

2000

CASE1, DABCASE2, spary combustionCASE2, rocket rich fuel

Ma=1.8 , H=10km

Rocket pressure =1.5Mpa

Rocket pressure =4.0Mpa

19%25%

Fig.7 With 2nd fuel injection, the relationship between fuel mass flow rate and thrust

2nd fuel mass flow rate/ (Kg/s)

Eje

ctio

nm

ass

flow

rate

/(K

g/s)

0 0.05 0.1 0.15 0.21.4

1.6

1.8

2

CASE1, DABCASE2, spary combustionCASE2, rocket rich fuel

Rocket pressure =4.0Mpa

Rocket pressure =1.5Mpa

Ma=0.8 , H=0km

8%

5%

2nd fuel mass flow rate/ (Kg/s)

Eje

ctio

nm

ass

flow

rate

/(K

g/s)

0 0.05 0.1 0.15 0.21

1.1

1.2

1.3

1.4

CASE1, DABCASE2, spary combustionCASE2, rocket rich fuel

Ma=1.8 , H=10km

9%

4.5%

Rocket pressure =1.5MpaRocket pressure =4.0Mpa

Fig.8 With 2nd fuel injection, the relationship between fuel mass flow rate and ejection mass flow rate

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O2.Mass_Fraction: 0.02 0.06 0.1 0.14 0.18 0.22

Fig.9 Distribution of O2 in combustion chamber with multistage rocket ejection and 2nd fuel added

From Fig. 7 to 9, we can see, secondary combustion organized through kerosene spray, may further improves

the utilization of ejected air, thus enhances the thrust force, but simultaneously actually reduces the ejector effect. Take the flight Mach number of 0.8, primary rocket chamber pressure of 4.0Mpa as an example, while realizes the thrust force to promote 15%, the ejection mass flow rate drops 5%.

C. 1st stage rocket of equivalent status, 2nd stage rocket of fuel-rich with 2nd rocket chamber pressure increase The secondary combustion is modeled under the conditions that 1st stage rocket of equivalent status, and the 2nd

stage rocket of equivalence ratio 1.5. Through increased 2nd rocket chamber pressure to increase fuel mass flow rates.

Thrust and the ejection mass flow rate changing with the secondary fuel mass flow rate are also seen in Fig. 7 and Fig. 8.

From Fig. 7 and Fig. 8, we can see that by raising 2nd stage fuel-rich rocket chamber pressure for secondary combustion, may simultaneously realize the thrust force enhancement and ejection mass increase. Take the flight Mach number of 1.8, rocket chamber pressure of 4.0Mpa as an example, while realizes the thrust force to promote 19%, the ejection mass flow rate promotes 9%.

Regarding the two methods of secondary combustion organization in the multistage rockets structure, from the view of gain largest thrust, kerosene spray combustion can achieve the better performance, but with the expense of loss of some ejection mass flow rate. Take the largest amount of ejection mass flow rate as the standard, improve the fuel-rich 2nd rocket chamber pressure, could enhance the thrust as well as increase the ejection mass. In fact, the increase of ejection mass flow rate means the decrease of resistance from the inlet of engine. Considering the integration of thrust and the resistance of whole engine, it is suggested that raising the fuel-rich 2nd rocket chamber pressure to organize secondary combustion should be the best choice.

Previous studies showed that in DAB mode of combustion, in the actual experiments, the secondary fuel injection location will greatly affect the engine performance and flame stability. In SMC mode of combustion, although easy to organize, but takes the price of actually large decline of combustion efficiency. In SPI mode, although carried on DAB and SMC the integration, but the overall performance is actually between the two. The author believes that no matter in any mode of secondary combustion, both the DAB mode, SMC mode and SPI mode, the chamber increased pressure "without blocking" affect the inlet of engine, resulting in ejection mass flow rate decreases. The increase of thrust and reduce the amount of ejection mass comes to be a pair of sharp contradictions.

The author believes that, for multistage rocket ejection proposed in this paper, the 1st stage rocket ejects incoming air, simultaneously plays the role of anti-chamber-pressure at the same time. The 2nd stage rocket could eject the flow in combustion chamber continuously, that is to say playing a "relay" ejector effect, and works as the flame holder simultaneity, while the fuel-rich flame plume works as a reliable promise of flame stabilization. The mechanism of multistage rockets ejection is the optimized combination of DAB and SMC, and integrates the advantages of the two. Not only may raise the efficiency of secondary combustion, but may also suppress the chamber pressure infecting the inlet, increase the ejection ratio. It may be the better choice under current research of the RBCC engine in ejector mode for secondary combustion.

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V. Conclusion Detail computational analysis of multistage rockets ejection has been performed in the present study. Some

conclusions from numerical simulation are: 1) Multistage rockets ejection show better performance than ordinary rockets structure. At the state of 1st stage

rocket working at equivalent ratio and the 2nd stage rocket of fuel-rich, the thrust of the whole engine could be increased associated with ejection mass flow rate also increased.

2) In multistage rockets, the 1st stage rocket not only ejects the air coming from far-field, but also suppresses the chamber pressure infecting the inlet. The 2nd stage rocket could eject the flow in chamber continuously, and works as the flame holder simultaneity.

3) The mechanism of multistage rockets ejection is the optimized combination of DAB and SMC. It may be the better choice under current research of the RBCC engine in ejector mode for secondary combustion.

References 1 A.Siebenhaar, M.J. Bulman and D.K. Bonnar, “The Role of the Strutjet Engine in New Global and Space Markets,” IAF-

98-S.5.04. 2 Lehman, M., Pal, S. and Santoro, R. J., “Experimental Investigation of the RBCC Rocket-Ejector Mode,” AIAA Paper

2000-3725. 3 Landrum, D. B., Thames, M., Parkinson, D. and Gautney, S., “Investigation of the Rocket Induced Flow Field in a

Rectangular Duct,” AIAA 1999-2100. 4 S.Muller, Clark.W.Hawk, P.G.Bakker, “Mixing of Supersonic Jets in a RBCC Strutjet Propulsion System,” AIAA Paper

99-2973 5 WANG Guo-hui, “Operation process research on ejector mode of rocket based combined cycle(RBCC),” Ph.D Dissertation,

Graduate school of Northwestern Polytechnical University, 2002.12.

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