Innovative concept proposal Space for Innovation;
An Electrical turbojet Engine.
Wijnand Schoemakers
Msc Aerodynamics 2013
My concept is a fully electrical turbojet engine. In order to fly a fully electrical plane at high
transonic mach numbers a turbojet engine can be preferred above an open rotor/ducted fan electrical
engine, because of the larger rotors reaching supersonic speeds.
The layout will be similar to a conventional turbojet engine, with the exception of the turbine.
The electrical engine layout will include an inlet duct, a fan and several compressor stages driven by a
generator, an electrical heater instead of a combustion chamber and a nozzle. This concept can be
analyzed using standard jet engine theory.
The main advantages for an electrical engine will be a (potentially) lighter weight because of
increased overall efficiency, and the removal of the turbine. This will be balanced by the additional mass
of the generator and the electrical heater. The engine will be more efficient because the generators will
be more efficient than the old turbine/shaft combination. The maximum heater temperature can be
much higher than the combustion temperature in a conventional engine because there is no turbine,
therefore also increasing the overall efficiency.
Two major bottlenecks can be identified with this concept; the energy density of batteries, and
the efficiency of the heater. For an electrical turbojet engine to be competitive with conventional
turbojet engines, the energy density of batteries has to increase by a lot. Batteries will have to be able to
compete not only with the energy density of the conventional fuels, but later also with hydrogen. I
believe this higher energy density is possible by 2040 due to the continuous increase spending in the
development of batteries in other industries, like the car industry. The advantage of this bottleneck is
that the energy density of batteries will continue to increase for probably a long time. Energy density of
hydrogen will almost remain constant on the other hand. This effect will result in continuously lighter
electrical systems, giving greater advantage over fossil/hydrogen fueled planes.
Secondly, an efficient heater needs to be developed that can transfer large amount of heat to
moving air without major losses in pressure. This will be a field of research mainly dependent on the
aerospace industry itself.
An electrical turbojet engine can be very sustainable for several reasons. Firstly, it will produce
zero emission during flight. Secondly, the increased efficiency over conventional jet engines will result in
less energy used. If the electrical turbojet engine is combined with other innovative features of the
aircraft this sustainability can even further be increased. For example, the airplane structure could be
integrated with low cost, light weight solar panels to generate inflight power. Also, during low speed
flight the heat exchanger could be turned off for increased efficiency and it could turn on at higher flight
speeds.
In order to determine the advantages of the electrical turbojet engine, I wrote a small matlab
script comparing the performance of a GE90-94B turbofan engine with a fully electrical engine. A lot of
assumptions were made during the calculations of the electrical engine. Therefore, I could not determine
the overall fuel efficiency of an electrical engine. However, the electrical engine did produce a lot higher
thrust and specific thrust than the GE90 engine at equal heater/combustion exit temperatures. This
would already suggest a better performance. Results were electrical thrust = 113 kN, conventional thrust
= 78.9 kN for equal mass flows.
In conclusion, I propose a fully electrical turbojet engine, using a generator to drive
fan/compressor and a heater. Major bottlenecks include battery technology and electrical heater
technology. However, when these bottlenecks are overcome, the electrical turbojet engine will be a
sustainable and efficient alternative to conventional turbojet engines. If selected for further review, a
better performance analysis will be done, and the potential bottlenecks will be further researched.
Appendix A – Matlab script
%% Turbofan Engine
% (GE90-94B, Twin Spool turbofan Engine)
% Wijnand Schoemakers
% 13-01-2012
clc,clear,close
% Input parameters in SI-units
Ta=216;
Pa=22632;
Ma=.8;
alt=11000;
eta_in=0.93;
pi_fan=1.65;
alpha=8.2; % Bypass ratio
pi_lpc=1.4;
pi_hpc=19;
To4=1550;
eta_fan=0.9;
eta_lpc=0.88;
eta_hpc=0.88;
eta_lpt=0.92;
eta_hpt=0.92;
eta_shaft=0.99;
eta_b=0.99;
% Burner efficiency
pi_b=0.96;
eta_n=0.97;
R=287;
LHV=43e6;
cp_air=1000;
gamma_air=1.4;
cp_gas=1150;
gamma_gas=1.33;
cor_m=1400;
aa=sqrt(gamma_air*R*Ta);
ua=Ma*aa;
% Inlet 1-2
To2=Ta*(1+(gamma_air-1)/2*Ma^2);
Toa=To2;
Poa=((1+(gamma_air-1)/2*Ma^2))^(gamma_air/(gamma_air-1))*Pa;
Po2=((1+eta_in*(gamma_air-1)/2*Ma^2))^(gamma_air/(gamma_air1))*Pa;
% Mass flow
delta=Poa/101325;
theta=Toa/288.15;
me=(cor_m/sqrt(theta)*delta);
m=me/(alpha+1);
% Engine core flow
ms=m*alpha;
% Engine secondary flow
% Fan 2-7
Po7=pi_fan*Po2;
To7=To2*(((pi_fan)^((gamma_air-1)/gamma_air)-1)/eta_fan+1);
% Compressor 2-3
Po3lpc=pi_lpc*Po7;
To3lpc=To7*((pi_lpc^((gamma_air-1)/gamma_air)-1)/eta_lpc+1);
Po3hpc=pi_hpc*Po3lpc;
To3hpc=To3lpc*((pi_hpc^((gamma_air-1)/gamma_air)-1)/eta_hpc+1);
% Burner 3-4
Po4=Po3hpc*pi_b;
mf=m*cp_gas*(To4-To3hpc)/(LHV*eta_b);
f=mf/m;
mt=mf+m;
% Turbine 4-5
To5hpt=To4-m*cp_air*(To3hpc-To3lpc)/eta_shaft/mt/cp_gas;
Po5hpt=Po4*(1-(1-(To5hpt/To4))/eta_hpt)^(gamma_gas/(gamma_gas1));
To5lpt=To5hpt-(m*cp_air*(To3lpc-To2)+alpha*m*cp_air*(To7To2))/eta_shaft/mt/cp_gas;
Po5lpt=Po5hpt*(1-(1(To5lpt/To5hpt))/eta_lpt)^(gamma_gas/(gamma_gas-1));
To5=To5lpt;
Po5=Po5lpt;
% Primary Nozzle 5-8
Ps8=Po5*(1+(1gamma_gas)/(eta_n*(1+gamma_gas)))^(gamma_gas/(gamma_gas-1));
if Pa>Ps8
disp('Primary Nozzle is unchocked')
else
P8=Ps8;
T8=2*To5/(1+gamma_gas);
u8=sqrt(gamma_gas*R*T8);
disp('Primary Nozzle is chocked')
end
% Fan nozzle 7-9
Ps9=Po7*(1+(1gamma_air)/(eta_n*(1+gamma_air)))^(gamma_air/(gamma_air-1));
if Pa>Ps9
disp('Fan Nozzle is unchocked')
else
P9=Ps9;
T9=2*To7/(1+gamma_air);
u9=sqrt(gamma_air*R*T9);
disp('Fan Nozzle is chocked')
end
rho8=P8/(R*T8);
rho9=P9/(R*T9);
A8=mt/(rho8*u8);
A9=ms/(rho9*u9);
% Thrust
F=mt*u8-me*ua+A8*(P8-Pa)+ms*u9+A9*(P9-Pa)
SF=F/me
TSFC=mf/F*1000000
%% Electrical Turbofan Engine
% Wijnand Schoemakers
% 12-01-2012
clc,clear,close
% Input parameters in SI-units
% Conditions
Ta=216;
Pa=22632;
Ma=.8;
alt=11000;
% Engine characteristics
To4=1700; % Heater final temperature
alpha=8.2; % Bypass ratio
cor_m=1400; % Corrected mass flow
% Efficiencies
eta_in=0.93;
% Inlet
eta_fan=0.9;
% Fan
eta_lpc=0.88;
% Low pressure compressor
eta_hpc=0.88;
% High pressure compressor
eta_n=0.97;
% Nozzle
% Pressure ratios
pi_fan=1.65;
%
pi_lpc=1.4;
%
pi_hpc=19;
%
pi_h=0.9;
%
Fan
Low pressure compressor
High pressure compressor
Heater
% Constants
R=287;
LHV=43e6;
cp_air=1000;
gamma_air=1.4;
% Perfomance calculation
aa=sqrt(gamma_air*R*Ta);
ua=Ma*aa;
% Speed of sound
% Inlet 1-2
To2=Ta*(1+(gamma_air-1)/2*Ma^2);
Toa=To2;
Poa=((1+(gamma_air-1)/2*Ma^2))^(gamma_air/(gamma_air-1))*Pa;
Po2=((1+eta_in*(gamma_air-1)/2*Ma^2))^(gamma_air/(gamma_air1))*Pa;
% Mass flow
delta=Poa/101325;
theta=Toa/288.15;
me=(cor_m/sqrt(theta)*delta);
m=me/(alpha+1);
% Engine core flow
ms=m*alpha;
% Engine secondary flow
mt=m;
mf=0;
% Fan 2-7
Po7=pi_fan*Po2;
To7=To2*(((pi_fan)^((gamma_air-1)/gamma_air)-1)/eta_fan+1);
% Compressor 2-3
Po3lpc=pi_lpc*Po7;
To3lpc=To7*((pi_lpc^((gamma_air-1)/gamma_air)-1)/eta_lpc+1);
Po3hpc=pi_hpc*Po3lpc;
To3hpc=To3lpc*((pi_hpc^((gamma_air-1)/gamma_air)-1)/eta_hpc+1);
% heater 3-5
Po5=Po3hpc*pi_h;
To5=To4;
% Primary Nozzle 5-8
Ps8=Po5*(1+(1gamma_air)/(eta_n*(1+gamma_air)))^(gamma_air/(gamma_air-1));
if Pa>Ps8
disp('Primary Nozzle is unchocked')
else
P8=Ps8;
T8=2*To5/(1+gamma_air);
u8=sqrt(gamma_air*R*T8);
disp('Primary Nozzle is chocked')
end
% Fan nozzle 7-9
Ps9=Po7*(1+(1gamma_air)/(eta_n*(1+gamma_air)))^(gamma_air/(gamma_air-1));
if Pa>Ps9
disp('Fan Nozzle is unchocked')
else
P9=Ps9;
T9=2*To7/(1+gamma_air);
u9=sqrt(gamma_air*R*T9);
disp('Fan Nozzle is chocked')
end
rho8=P8/(R*T8);
rho9=P9/(R*T9);
A8=mt/(rho8*u8);
A9=ms/(rho9*u9);
% Thrust
F=mt*u8-me*ua+A8*(P8-Pa)+ms*u9+A9*(P9-Pa)
SF=F/me
TSFC=mf/F*1000000