Presented at Infotech@aerospace2011 AIAA Paper 2011-1634

1. Infotec@Aerospace 2011
I@A-89 Evolutionary Design of Intelligent Systems AIAA-2011-1634
Multidisciplinary and Multi-objective Design Exploration
Methodology for Conceptual Design of a Hybrid Rocket
Yukihiro Kosugi
Tokyo Metropolitan University (TMU)
Akira Oyama
Japan Aerospace Exploration Agency
Kozo Fujii
Japan Aerospace Exploration Agency
Masahiro Kanazaki
Tokyo Metropolitan University (TMU)

2. Contents 2
 Background
 Objectives
 Design method
 Evaluation procedure of LV with HRE
 Multi-objective Genetic Algorithm (MOGA)
 Scatter Matrix Plot (SPM)
 Design problem for LV with HRE
 Design variables
 Objective functions
 Results
 Design results
 Visualization of non-dominated solutions
 Design knowledge
 Conclusions

3. Background1 Rockets presently used for space transportation 3
 Solid-propellant rocket engine
Advantage：・Simple mechanism and construction
・Easy to maintain the propellant
Disadvantage：・Low specific impulse (Isp)
・Inability to stop combustion after it is ignited
・Environment issues
(caused by ammonium perchlorate (NH4ClO4),
and aluminum oxide (Al2O3))
 Liquid-propellant rocket engine
Advantage ：・High specific impulse (Isp)
・Ability to stop/restart combustion
Disadvantage：・Complex mechanism and construction
・Risk of explosion
・Difficulty to store low temperature propellant

4. Background2 What is hybrid rocket? 4
Hybrid Rocket Engine(HRE) ： propellant stored in two kinds of phases
It can adopt the beneficial features of both the liquid and solid rockets.
Solid fuel + Liquid oxidizer :
Advantage of HRE
・Simple construction and mechanism
・Higher specific impulse (ISP) than solid rocket engine
・Ability to stop/restart combustion
・Low environmental impact and low cost

5. Background4 Design of HRE 5
 Solid rocket：Premixing type solid propellant
 Fuel rocket：Mass flow control of fluid propellant
→ Easy to maintain a constant oxidizer and mass -
fuel ratio (O/F) and to get a stable thrust
 HRE：The mixture of fuel and oxidizer is initiated after ignition.
Combustion occurs in the boundary layer diffusion flame.
→ Because O/F is decided in this part of combustion process, the solid fuel
geometry and the supply control of the oxidizer have to be optimally
combined.
⇔With excessive mass flow of oxidizer, the
rocket achieves higher thrust, but structural
weight should be heavier .
Importance to find optimum fuel geometry and oxidizer supply
⇒Multi-disciplinary design which is considered propulsion,
structure and trajectory

6. Background５ Meta-heuristics approach in aircraft design 6
Evolutionary Algorithm based Design Exploration
Application of Mitsubishi Regional Jet (MRJ)
Targets
•Wing design
•High-lift Airfoil design
•Nacelle chine design
Design Exploration
•Genetic Algorithm
•Surrogate model
•Data mining
・Chiba, K., Obayashi, S., Nakahashi, K., and Morino, H., "High-Fidelity Multidisciplinary Design Optimization of Aerostructural Wing
Shape for Regional Jet," AIAA Paper 2005-5080, AIAA 23rd Applied Aerodynamics Conference, Toronto, Canada, June 2005.
・Kanazaki, M., and Jeong, S., “High-lift Airfoil Design Using Kriging based MOGA and Data Mining,” The Korean Society for
Aeronautical & Space Sciences International Journal, Vol. 8, No. 2, pp. 28-36, November 2007.
・Kanazaki, M., Yokokawa, Y., Murayama, M., Ito, T., Jeong, S., and Yamamoto, K., “Nacelle Chine Installation Based on Wind Tunnel
Test Using Efficient Design Exploration,” Transaction of Japan Society and Space Science, Vol.51, No. 173, pp. 146-150, November
2008. … etc.
Design Exploration is also expected in MDO for hybrid rocket.

7. Objectives of this study 7
Development of the evaluation tool for concept of
launch vehicle (LV) with HRE
Evaluation based on the empirical model
Demonstration of multi-disciplinary design using
genetic algorithm
Conceptual design of single stage surrounding
rocket which achieves low gross weight and high
flight altitude
Knowledge discovery using data mining

8. Design method1 evaluation1 8
Overview of the evaluation procedure
Input variable Output variable
* Mass flow of oxidizer [kg/s] * Flight altitude [km]
* Fuel length [m] * Gross weight [kg]
* Port radius of fuel [m] * Total oxidizer weight [kg]
* Combustion time [s] * Total fuel weight [kg]
* Pressure of combustion chamber [MPa] * Nozzle length [m]
* aperture ratio of nozzle [-] * Combustion chamber length [m]
* Oxidizer tank length [m]
* Rocket radius [m]
* Rocket aspect ratio [-]
* Nozzle throat area [m2]
* Thrust at ignition [kN]
* Initial oxidizer mass flux [kg/m2s]
* History of flight, thrust, and
combustion chamber pressure

9. Design method2 evaluation2 9
O/F and thrust power estimations
Underlined variables are part of the design parameters.
moxi t 

O F (t ) 
m fuel t 

rport (t )  a  Goxi t 
 n
Mass of vaporized fuel: m fuel t   2rport t L fuel  fuel rport (t )
 
A
NASA-CEA
t
rport (t )  rport (0)   rport ( )d

0

T t   T    m propue  Pe  Pa  Ae
 
m prop (t )  moxi (t )  m fuel (t )
   fuel
t
moxi t 
 = mass flow of oxidizer m prop t 
 = mass flow of propellant rport (t )  rport (0)   rport ( )d

m fuel t 
 = mass flow of fuel 0
Peh = pressure in the combustion chamber
Lfuel = length of fuel Pe = pressure at nozzle exit
ρfuel = density of fuel (constant) Pa = pressure of atmosphere at flight altitude
rport = radius of fuel port ue = velocity at nozzle exit
T(t) = thrust ηT λ = momentum loss coefficient at nozzle exit by friction (<1)
= total thrust loss coefficient by deflection of propellant at nozzle exit (<1)

10. Design method3 evaluation3 10
Weight and length estimations
Mtot is estimated by the sum of the components’ weight
moxi t dt
tburn
M tot  M engine  M pay  M ex M oxi   
0
m fuel t dt
tburn
M engine  M oxi  M fuel  M res  M ch M fuel   
0
2
M ex  M engine M res  Vres
3
Ltot is one and a half times as long as HRE for including payload. M ch  Vch
tburn = combustion time
Moxi = mass of total oxidizer
Mfuel = mass of total fuel
Mres = mass of the oxidizer tank
Mch = mass of the combustion chamber The sketch of the oxidizer tanks. The sketch of the nozzles.
Vres = integrated volume of material for the oxidizer tank
Vch = integrated volume of material for the combustion chamber
Mtot = gross weigh Ltot = total length of the rocket
Mengine = engine weight Lch = length of the combustion chamber (=Lfuel)
Mpay = payload weight Lres = length of the oxidizer tank
Mex = weight of other equipments Lnozzle = length of the nozzle

11. Design method4 evaluation4 11
Trajectory prediction S-520: JAXA’s surrounding solid propellant rocket
Equation of motion
T t   Dt 
at   g
M tot t 
D(t) = total drag at time t
Friction drag coef. of outer surface of the rocket
A(t) = acceleration at time t
0.455
C D f , Design (t )  g = gravitational acceleration
log 10 Re2.58 1  0.144M 
2 0.65 Dp, Design(t)
Df, Design(t)
=
=
pressure drag at time t
friction drag at time t
Pressure drag coeff. based on S-520 flight data Re = Reynolds number
S wet ,S 520 CDp, Design(t) = pressure drag coefficient at time t
C D p , Design (t )  C D,S 520 (t )  C D f ,S 520 (t ) CDf, Design(t) = friction coefficient drag at time t
S ref ,S 520 CDp, S-520(t) = pressure drag coefficient of the solid rocket S-520
 D p , Design (t )  qS ref , DesignC D p ,S 520
CDf, S-520(t) = friction drag coefficient of the solid rocket S-520 a

D
q = dynamic pressure
 f , Design (t )  qS wet , DesignC D f , Design

Sref, Design = area of cross section of the designed rocket
Swet Design = wetted area of cylinder of the designed rocket
D(t )  D p , Design (t )  D f , Design (t ) Sref, S-520 = area of cross section of the solid rocket S-520
Swet S-520 = wetted area of cylinder of the solid rocket S-520
The aerodynamic effect of the rocket length and diameter can be evaluated.

12. Design method5 design method1 12
 Heuristic search：Multi-objective genetic algorithm
(MOGA)
 Inspired by evolution of life
 Selection, crossover, mutation
 Global search
 Pareto-ranking method
 Ranking of designs for multi-objective function
Sub-population
Individual
Individual
 Island model
Migration
Sub-population
(a) (b)
Hiroyasu, T., Miki, M. and Watanabe, S., “The New Model of Parallel
Genetic Algorithm in Multi-Objective Optimization Problems (Divided Range
Multi-Objective Genetic Algorithm),” IEEE Proceedings of the Congress on
Evolutionary Computation 2000, Vol. 1, pp. 333-340, 2000.

13. Design method6 design method2 13
Scatter Plot Matrix (SPM)
→For the design problem investigation
a Scatter plot of avs.b
b
Correlation of avs.b
c
d
SPM arranges two-dimensional scatter plots among attribute values like a matrix
・The present SPM shows scatter plots on the upper triangular, and
correlation coefficient on the lower triangular (Software R is used for statistical
computing and graphics.)

14. Design problem for LV with HRE1 14
Swirling oxidizer type HRE
 Proposed by Yuasa, et al.
 Swirling oxidizer is supplied into the fuel.
 Polypropylene is employed as a fuel.
rport t   0.0826Goxi
This expression was
 0.55 provided by Prof. Yuasa.
regression rate against the mass flux of the oxidizer
Yuasa, S., et al, “Fuel Regression Rate Behavior in Swirling-Oxidizer-Flow-Type Hybrid Rocket Engines,” Proc 8th International
Symposium on Special Topics in Chemical Propulsion, No. 143, 2009.

15. Design problem for LV with HRE 15
Design target
the surrounding rocket with assuming that the 40kg payload is carried
Design variables (6)
Lower Upper
Mass flow of oxidizer [kg/s](dv1) 1.0 30.0
Fuel length [m] (dv2) 1.0 10.0
Port radius of fuel [m] (dv3) 0.01 0.30
Combustion time [s] (dv4) 10.0 40.0
Pressure of combustion chamber [MPa] (dv5) 3.0 6.0
aperture ratio of nozzle [-](dv6) 5.0 8.0
Objective functions (2)
Minimize Gross weight, Wgross
Maximize Maximum flight altitude Hmax

16. Results1 Sampling results by MOGA
16
 MOGA result colored by rocket’s aspect ratio (length/diameter)
After 100 generation started with 64 individuals
Non-dominated solutions
The solutions for which Hmax is greater
than 150 km have a larger Wgross than the
solutions for which Hmax is less than 150
km.
To achieve high flight altitude, the rocket’s
aspect ratio becomes high.
Optimum direction
-There is trade-off between Wgross and Hmax.
-Maximum Hmax is about 180km.
- The rocket considered here is suitable for the sub-orbital flight around 100km altitude.

17. Results2 Comparison with JAXA’s S-210
17
S-210 Des1
Flight altitude [km](H) 110.0 108.2
Gross weight [kg](W) 260.0 334.0
Payload weight [kg] 40.0 40.0
Payload weight/gross weight 0.154 0.120
Fuel length [m] 5.2 8.5
Fuel diameter [mm] 210.0 210.0
Rocket’s aspect ratio [-] 24.8 40.5
-Comparison with sub-orbital solid rocket S-210
which flight about altitude 100km
-Aspect ratio of the HRE rocket was larger than that
of the S-210, because the oxidizer tank and the solid
fuel are placed longitudinally in the rocket.
-Wpey/Wgross of the designed rocket was less than
www.isas,jaxa,jp
that of the S-210. → Pch is supposed to be constant
during the combustion process

18. Results3 Visualization of non-dominated solution by SPM
18
dv3(port diameter in the fuel) of non-
dominated solutions becomes less.
→ better volumetric efficiency and
slender chamber
There are correlation among dv1(mass
flow of oxidizer), dv2(fuel length), and two
objective functions.
The rockets’ aspect ratio and the Acc_max
are correlative relation.

19. Results4 Design knowledge from non-dominated solution
19
 There is a trade-off between the minimization of the gross
weight and the maximization of the flight altitude.
 The designed rockets achieve a higher flight altitude by
reducing the aerodynamic drag and by employing a higher
aspect ratio.
 The rockets which is lower oxidizer mass flow, the
diameter of the combustion chamber becomes smaller. As
this result, the weight of the combustion chamber
decreases.
 The rocket with HRE, which achieves a higher flight
altitude and lower weight, tends to be long and narrow in
the present MDO problem

20. Conclusions 20
 Development of the design tool for concept of LV) with HRE
 Evaluation based on the empirical model
 Evaluation module (cygwin script) is open to the public,
http://bit.ly/hmBYgB.
 English manual will be also uploaded in this April-May. This
uploading will announced in twitter, (@tmu_craft_desig #hre).
 Knowledge discovery of multi-disciplinary design for a LV with
HRE
 High aspect ratio rocket is better for the present design problem.
 With heavier mass of the oxidizer, the rocket’s gross weight becomes
heavier.
 Future work:
 The sophisticate of the evaluation tool.
 Consideration of the pressure time variation of the combustion chamber.
 Structural type for realistic rocket.
 Conceptual design of orbital rocket for microsat launch

21. Acknowledgement 21
We thank members of the hybrid rocket engine
research working group in ISAS/JAXA for giving
their experimental data and their valuable advices.
This paper and presentation was supported by
ISAS/JAXA.
Thank you very much for your kind attention.