Effect of Inert Propellant Injection - NASA Technical Reports Server ...

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AIAA,92-3447

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Effect of Inert Propellant Injection- on MarS Ascent Vehi¢l_ Performance ....... i.....

James .............

189238._

E. Colvin

_i

........

University of Arizona Space Engineering Research Center Ttieson, Arizona

- - -L _ 77 _ =................................... ....

and Geoffrey A. Landis Sverdrup Technolog); Inc. Lewis Research Center_Group Brook Park, Ohio = ........

Prepared for the 28th Joint Propulsion cosponsored Nashville,

Conference

by the AIAA, Tennessee,

IW A

_ _ ____

_

--- ' :

and Exhibit

==

SAE, AS_M_E, and ASEE

July 6-8,

1992

(NASA-CR-189238) EFFECT PQ0PELLANT INJECTION qN VEHICLE (Sverdrup

PER_GRMANCE Technology)

OF MARS

Final 30

INERT ASCENT

N93-12152

Report p

Unclas

G3/18

0123621

...

EFFECT OF INERT MARS ASCENT

PROPELLANT INJECTION VEHICLE PERFORMANCE

James

ON

E. Colvin

University

of Arizona

Space

Engineering

Research

Center

Tucson,

Arizona and

Geoffrey

A. Landis

Sverdrup Technology, Lewis Research Center Brook

Park,

Inc. Group

Ohio

ABSTRACT

A Mars

ascent

vehicle

from Earth.

In some

the engine,

if the inert

from

Earth.

which

Carbon

could

cases

is limited

in performance

the vehicle

performance

gas is available dioxide,

be separated

as an in-situ

nitrogen,

and argon

by compressing

by the propellant can be improved resource

can be brought

by injecting

and does

are constituents

the atmosphere,

which

not have

inert

gas into

to be brought

of the Martian

atmosphere

any chemical

processing

without

step. The

effect

numerical

combustion

combustion

chamber

Results

of this

required

mass

propellant

of inert

code pressure,

analysis of return

injection.

gas

were

injection that

on rocket

calculated

expansion applied

propellant

engine

chemical

ratio,

equilibrium

oxidizer/fuel

to several

needed

performance

ratio,

candidate

in low Earth

orbit

analyzed

for engines and inert injection

missions could

was

to determine be decreased

with

a

of varying fraction. how using

the inert

NOMENCLATURE

AR

Area ratio of a nozzle

cIa4-o2

methane/oxygen

CO-O 2

carbon-monoxide/oxygen

F

thrust

g

gravitational

H2-O 2

hydrogen/oxygen

engine fueled

engine

(nt) constant

fuel+oxidizer

Isp(FO)

fueled

Isp(P)

propellant

Mc

Earth

of Earth,

fueled specific

specific

return

engine

impulse

impulse

payload

9.82 rn/sec.

(kg),

(sec);

(sec);

thrust per unit fuel and oxidizer

thrust per unit propellant

including

crew

capsule,

crew,

mass

flow

mass flow suits,

consumables,

and soil samples. ME

engine

rhF+ O

mass flowof

and oxidizer

Mf

without

mass (kg) the energy-producing

components

of the propellant

(kg/sec)

the inert).

burnout

mass of stage (kg)

initial mass of stage (kg)

lVlp

propellant

rhtotal

total mass flow of propellant

mass (kg)

tank and structure

mass

(kg/sec)

(kg)

N2

nitrogen

O/F

ratio of mass flow rate of oxidizer

AV

velocity

increment,

to mass

km/sec

2

flow rate of fuel into engine

(fuel

iI

INTRODUCTION It is possible if Mars

derived

return Mars

for a return resources

to Earth.

The

for the return

production every

1"5.

landing

Most

locally

word

extremely

a factor simple:

Table

I shows for inert

to designate

mass

a gas

production

requires

the composition gas available

combustion which

adds

of fuel

While or oxidizer,

only compressing of the Martian

from

the Mars

mass

atmosphere

to

is present

at

gas

separation

and

it is worthwhile

processing. Earth

would

of the return

be to inject

engines.

to the exhaust

(Here

but does

not

this does not decrease

fuel mass

it has the advantage

of being

and liquefying

atmosphere.

(LEO)

to acquire.

However,

from

flow

for the

in propellant

which

involve

no chemical

brought

with the propellant.)

as chemical the process

require

orbit

resources

operations

of Mars

less costly

required

low Earth

atmosphere,

fuel and oxidizer.

which

the propellant

reactions

or mining

of the atmosphere

gas into the propellant

"inert"

from

is the Martian

into rocket

significantly

or all of the propellant

by the use of local

prospecting

atmosphere

to reduce

"inert"

in chemical

by as great

choices

way

reduced

available

processing

uses of the Martian

the

participate

resource

some

to be made

that must be brought

is significantly

for utilization

chemical

available

we use

flight

site and does not require

possible

of Mars

can be used to provide

The simplest

energy-intensive

One

to the surface

mass of propellant

suggestions

to consider

mission

6

the atmosphere.

The

are carbon

most

easily

dioxide,

available

nitrogen,

and

energy

source

and

are separate

items.

In a

argon.

THEORY Rocket reaction

propellant

mass.

conventional

For some chemical

it is also possible is the chemical

rocket

rocket,

types,

fuel and oxidizer,

(and hence

case where

divided

source

reaction

mass

and the reaction the inert mass

the total impulse)

into

such as a nuclear

the energy

to add additional

In the idealized thrust

can be conceptually

increases

two

rocket,

and the reaction to the exhaust.

elements: these

mass are the same.

In this case the energy

mass is this plus the added

does not affect as the square

3

However,

the combustion

source

inert component. energy

root of the total mass

release,

flow:

the

F(wih in.t) =FO:+O_y)

m total _V

Specific

impulse

_

(Isp) is the thrust per unit of mass flow.

added into the propellant,

it is possible

additional

includes

inert mass.

inert reaction

mass is

two different ways.

We will

as the thrust per unit pro_oellant mass flow, where

both the energy producing The propellant

When

to define specific impulse

define the uropellant specific impulse Isp(P) the propellant

(1)

"

mr-+o

components

(fuel and oxidizer)

specific impulse thus decreases

plus the

as the square root of the

total mass flow: Isp(P)with inert injection = Isp(no into)

. ,V/_F,o

(2)

mto,,l

Here Isp(nO inert) is the specific impulse of the fuel/oxidizer rhtota 1is the total mass flow of propellant, components

of the propellant

system without inert injection,

and rhF+ O is the mass flow of the energy-producing

(fuel and oxidizer without the inert).

Alternately, we define the fuel+oxiclizer mass flow, not including

specific impulse lsp(FO) as the thrust per ut, it fuel

the additional

inert mass.

This is related to the propellant

specific impulse by:

Isp(F+O)= The fuel+oxidizer vehicle

mfud+m°x+min'_ . , . [ mox+miner/

Isp_)

specific impulse increases as the inert flow increase.

where the fuel and oxidizer

available,

the increase

decrease in propellant

must be shipped

in fuel+oxidizer

specific

maximize

of fuel and oxidizer

impulse

required

per ton of payload

delivered,

where

fixed mass such as tanks and engines is constant

than the

the payload.

during

This can be calculated

thrust,

This is equivalent

to the required mission delta-V:

to optimizing impulse

to the

may be

in closed form for the simple case

are ignored 7,8.-

the optimum

to minimize

or equivalently,

of a rocket where the energy source is fixed, but the specific

to maximize

proportional

may be more important

are brought from Earth, it is desired

varied

impulse

from Earth, but inert gasses are locally

the payload per unit amount of fuel and oxidizer.

performance

For a Mars-return

specific impulse.

For the case where the fuel and oxidizer the amount

(3)

-

is found

If the (propellant)

when

the specific

specific

impulse

is

glsp For

low

impulse,

mission

and

launch

adding

mass

= 0.625

AV

AV, the

increase

inert

required

(4)

reaction

by the

in thrust

mass

lower

outweighs

is effective.

specific

the loss

For high

impulse

means

of propellant

mission

that

specific

AV, the additional

adding

inert

mass

is not

effective. It is also worth example, velocity

noting

if the inert [glsp(t)]

that if the specific

mass

is exactly

impulse

flow

is allowed

equal

to the mission

is allowed

to vary),

to vary

the optimum

velocity.

during

occurs

This case was

the flight

when

(for

the exhaust

not examined

in this

study. The present (1) How

study

does

addressed

the addition

specific

impulse

reaction

are included

of rocket

There

engines

are two

Martian

reasons

for this. will

mass

not

to study

First, produce

and

complexity

to the

atmospheric

components.

Thus,

it is of

degradation

in performance

results

system The amount

content

second

compression

of the

consumption.

Thus,

energy-intensive To

study

from

the Martian

and

and chemical

atmosphere

of impulse

injection operating

of

100%

system

reduce

residual

the

and,

to remove

the

fact,

it may

unnecessary

the add

("inert") how

acceptable

and complexity

fuel. from

of Mars-derived

still produces

the mass

in

fuels

to quantify

composition

plant

in the fuel produced,

interest

by a fixed

Mars-derived

to produce

purity,

inert

produced

for entirely plant

considerable

production

is that propellant

available. Martian

Inert

gas,

atmosphere,

production on the

injection

using

Mars

on Mars other

a process

like the case of Earth-derived

fuels with inert injection, inert

the thrust

much fuel

or

propellant

if

of the separation

be reduced.

reason

of energy

an oxygen

remains

may possibly

from

inert

fuels

mass

50% nitrogen

affect

due to thermodynamics

the amount

an actual

considerable

If, for example,

derived

of increasing

it is also of interest

atmosphere

oxidizer.

loses

chamber

mission?

to the possibility

of fuel,

to the combustion

when real-world

of inert reaction

mass of a Mars return

In addition amount

of inert mass

in the analysis?

(2) Can the addition mission

two questions:

is likely

hand,

will

requiring

fuels,

it would

to be limited

be produced

by the

by simple

comparatively

low

energy

be desirable

to "stretch"

the

if this is possible resources,

5

a two

phase

parametric

study

was

undertaken.In phaseI, the computercodeComplexChemicalEquilibrium (CEC)9 was used to predict the effectson the vacuumspecific impulse (Isp) by varying the percentof inert loadinglevelsin

the total propellant

temperature,

pressure,

using

thermodynamic

known

In phase

this study

and chemical

injection

would

inert propellant

into the exhaust

propellant

combination

is typical

02) propellant to store, from

OF

Earth 3'4.

manufactured Likewise, nitrogen,

specific

the three

CEC,

mass

be 1.0 kg/sec. mixture gas

ratio

was,

oxidizer independent whether

the inert

inert

amounts

ROCKET

The

impulse

conclusion

be realized

of

from

by injecting

impulse,

EFFICIENCY

hydrogen/oxygen

system

available.

but the propellant methane

on Mars

(CO-O

gasses

(H2-O

fuel

2)

Methane/oxygen

(CH 4-

is more compact,

easier

if hydrogen

2) propellant

in the Martian

assumed

arbitrarily

of fuel

is brought

system

might

atmosphere,

carbon

the addition

of one

sets the total propellant

and oxidizer

to be added

Practical

is a variable.

to the mass in the same

the combustion

injection. added

by investigating

flow rate of oxidizer

this allows

is best

The desired

the effect

level to 0, 25, 50, 75, 90, and 95 percent

The CEC code

relative

inert

scenarios.

ON

studied.

was examined

or the mass

of the

used to examine

monoxide/oxygen

the inert loading

flow-rate.

the fuel;

products.

CEC were

could

and nozzle

be

dioxide,

considered.

for convenience, and

chamber

resources.

each propellant

(O/F)

and exhaust

the equilibrium

engines.

of manufacturing

most common

were

The

specific

the carbon

time and then by varying propellant

lower

only Mars

and argon,

Using

of the highest

combustion

from LEO

INJECTION were

is a possibility

using

savings

flow of the return

combinations

Finally,

from mission

a mass

INERT

has somewhat

and there

of the propellant

proposed

whether

This code calculates

of the rocket

calculated

on several

EFFECT Three

impulses

be to show

flow-rate.

composition

parameters

II, the specific

inert propellant

combustion

to have

engine

with the oxidizer,

the same

design

the fuel,

flow-rate

proportion reaction

considerations or separately

to

is set as the

flow rate of fuel. mass

at a

of the total

mass This

inert

The

inert

to both

the

stoichiometry will determine

to the combustion

chamber. As an example

calculation,

H2-O 2 with OfF equaling

6.0.

consider

the case of adding

For the specified

propellant

6

25% inert to the baseline mass

flow-rate

propellant

of 1 kg/sec,

with no

inertthe0 2 flow-rate is 0.857kg/sec,and requires

0.2143

The flowrates a breakdown

kg/sec

were

of inert in the oxidizer

must then be reduced

of the propellant

flow-rate

composition

with the addition

from

Table

of fuel

II that the total

and oxidizer

inert

loading

Combustion

engine,

exit

however,

area

the

between

hydrogen,

the

in performance

200.

worse

nitrogen between

5 show

the

percentage

of inert

in the propellant

added

Similar

provided

input

PSI), representative

assumed there

the

calculations

of a pump-fed

that

with

for

of a

engine.

The

to operate

in vacuum,

is negligible

difference

surface. specific

fueled

impulse

engines. from carbon

almost

exactly

chosen

to be the inert similar

inert

injection

All of the

inert

dioxide

injection

the ideal

results

with

behavior

added.

behave

compared

(equation

Since

will apply

gasses

for

2).

there

to For

was little

for all.

fuels compared.

fuel+oxidizer

Isp mass

(as defined flow,

1, the fuel+oxidizer

above)

for the

specific

same

impulse

plotted

as a function

conditions

of the

as figure

increases

as inert

4.

As

propellant

is

to the mixture.

O/F mixture 0.55

is so low

inert gasses,

Figure

equation

was

were

Table II is

of the total,

calculations

1.38 MPa (200

engines

performance

follows

4 shows

from

The

flow.

of 25% inert gas.

is 25%

representative

in propellant

Figure

expected

PSI),

monoxide

behavior

the three

were

on Mars

change

carbon

of the study,

mixtures.

and on the Mars

with a slightly The

propellant

(3000

pressure

and

or argon.

the remainder

MPa

by 25%.

at an O/F of 6.0. of these

considered

was

in vacuum

1-3 shows

the same,

difference

(AR)

atmospheric

methane,

nitrogen

20.7

flow

The results

the three

pressures

and

ratio

operation

Figures

nearly

chamber

mass

still maintained

levels.

CEC and were used to investigate

inert

To add 25% inert

of inert in the hydrogen

components

used for other

nozzle

kg/sec

kg/sec.

propellant

amounts

pressure-fed

and 0.0357

0.143

of the original

It can be seen relative

the H 2 flow-rate

ratios

for CO-O 2 fueled

used for the previous engines.

These

figures

mixture

were:

ratios

6.0 for 1-I2-O 2, 3.4 for CH4-O2,

are nearly

optimal

for performance

and with

no inert injection. Similar figures mixture

results

6-11. ratios,

not depend

were

obtained

Here we show keeping

for the chamber

the effect

the AR constant

pressure

of the inert loading at 200.

of 20.7 MPa on Isp(P)

(3000

and IspfF+O)

It is clear that the optimum

on the inert injection.

7

PSI),

mixture

shown

in

for various ratio does

u

Finally, MPa

the effect

chamber

results

of inert

pressure

and

of this analysis.

loading O/F

is shown

close

for varying

nozzle

to stoichiometric.

The effectiveness

area ratio,

Figures

of inert injection

again

12 through

for 20.7

17 show

does not significantly

depend

the

on the

area ratio.

EFFECT The data from specific

impulse

of interest available

Phase

OF

INERT

I show

that

decreases,

two proposed

scenarios

direct launch from the surface in the "Mars

Direct"

was a one-stage These vehicle

were

propellant examined

specific

proposed

were chosen

analyses,

combinations engine

results

MISSION level

specific

parametrically.

were in less

increases.

The question

To answer

The first the mission The second

of the range

of values

of locally

this question,

was

a two-stage

parameters scenario

with an orbiting Earth return

as representative

used

analyzed

vehicle 10.

of Mars launch

for typical Mars return missions.

the inert used considered, advantage

the propellant

by the addition

impulse.'?

by Zubrin and Baker. 3,4

MASS

is increased,

impulse

to an Earth transfer orbit, following

and their AV requirements

For the mission

monoxide

loading

launch to Mars orbit for rendezvous

two scenarios

masses

propellant

mission

ON

of H 2 brought from LEO be reduced

in spite of the decreasing mission

as the inert

but the fuel plus oxidizer

is, can the amount inerts

INJECTION

was nitrogen. as the

obtained

considered.

8

lower

Only specific

the H2-O 2 and impulse

from inert injection

of the

CH4-O 2 carbon

for the mission

AVs

"MARS

DIRECT"

MISSION

1: ENGINE: For the purpose of the Advanced slightly.

The

at 89 kN.

Space Engine

assumed

In order

independent

12,250

two

was

stage

mission

on the

Mass

would

be very

and thrust

levels

The assumed same

basis,

thrust

similar

were modified

per engine

the thrust

was

to that

was fixed

assumed

to be

rate.

described

It was assumed

for each level

had the inert added

3: SECOND

by Zubrin

and

Baker

has

study

that the tank

for this parametric mass

per stage.

of inert injection

to its propellant

from

an Earth

The values

the rocket

return and

payload

structure

of mass

of Mp, M E, and M T were

equation

(5).

Only

the f'wst stage

mix.

STAGE:

The calculation as input

181 kg.

scenarios

be 10% of the propellant

determined

by Rocketdyne.

mass per engine to compare

that the engine

MASS:

kg (Mc).

(M T) would

mass

designed

of the inert injection

2: SYSTEM The

of this study, it was assumed

starts

with the second

to the calculation

burn was calculated

(upper)

for the first (bottom)

from the quoted

masses

using

stage,

and then

uses the second

stage calculation. the rocket

stage

AV for the second

total stage

equation:

AV = gIsp In [MM--_i 1

(5)

where: M i=M

e+Mp

MF=Mi-M Using

Zubrin

these

breakdown

(7) mass

values

the

is therefore

H2-O 2 was compared

for the second

AV of 3,350 m/see

of inert mass into the second

calculations

(6)

T

p

and Baker's

2,560 kg), a second-stage Injection

+M E+M

second

stage

only dependent

assumed

upon

(Mp

--!-22,170

kg and M E =

was calculated.

stage engine

was

stage

does not improve

to have

zero

inert

the type of propellant

with the CH4-O 2 propellant

assumed

by Zubrin

performance, injection.

used.

A baseline

and Baker.

and for The

mass

case

of

TableIH showsthe second-stage 495 sec.),

and Table

To verify final

the mass

that the assumed

acceleration

engine.

IV shows

acceleration.

thrust

of the second

Accelerations

mass

breakdown

levels

stage

are shown

The values

breakdown

using

burn

were

masses

assuming

V as a fraction

of Earth

tolerable

ranges

lisp =

lisp = 384 sec).

are reasonable,

calculated,

are well within

combination

CH4-O 2 propellant

and calculated

in Table

calculated

for H2-O 2 propellant

the initial

the use of one and Mars

and

89 kN

gravitational

for both hardware

and crew

safety. For the CH4-O 2 propellant, per kg of propellant.

The

the O/F ratio equals

total

amount

scenario,

only the hydrogen

is brought

chemical

processing

The relative

hydrogen

that must be brought

steps.

of CH 4 needed from Earth; masses

from LEO

In the case of the H2-O 2 propellant, of fuel in the propellant LEO

being

3.4, and thus 0.2273 is 4,724

methane

kg.

In the "Mars

is produced

of

of C and H in CH 4 is 3:1, and so the mass

of

for use in the second

The total amount

from

Direct"

it by a series

the O/F ratio equals

14.3%.

kg of CH 4 is needed

stage is 1,181 kg.

6.0, resulting

of hydrogen

in the mass

fraction

that must be brought

from

is 1,960 kg.

5. FIRST

STAGE

A AV of 3,400 the CH4-O 2 case stage

mass

ACCELERATIONS:

m/sec

was used for the fast

(no inert)

of 36,980

acceleration

was

limits.

15 engines

with a total mass

number calculated

of engines

loading

being

for two cases

levels

was

values

M E = 8,850

of 2,720

these

for the fast

kg, plus

stage for

the total

second

masses

to examine

first stage These

mission,

and the engines

and within these

engines

values

thrust throttled

to keep levels

VIII.

acceptable

at two extremes.

10

not exceed

would

as necessary).

of

for a total thrust

the number

of variables

be optimized

of to

with the

Accelerations

were

and 75% inert addition.

The acceleration ranges.

did

was the assumption

combine

mix: 0% inert addition

in Table

levels

calculations

were kept constant

(For an actual

are shown

for the H2-O 2 and CH4-O 2 case.

that the acceleration

kg (ME).

for each propellant

gravity

breakdown

to verify

throughout

adjusted,

accelerations

than one Mars

calculated

and engine

in this analysis.

Calculated greater

Consistent

The thrust

a minimum

kg and

the first stage mass

reasonable

1,330 kN.

= 70,160

Mass

kg.

Table VI and VII show Again,

are Mp

stage bum.

The purpose

found

is for all cases

for using

At an intermediate

these

two

loading,

the

accelerations

would

6. MASS

be between

these values.

OF HYDROGEN

The mass of hydrogen the rocket

equation,

propellant

mass):

REQUIRED

brought

modified

from LEO for the first stage can now be calculated.

to account

(_-1) Mp -

(Mc+

for tank and structure

fraction

(assumed

We use

to be 10% of

MO (10)

1.1-0.1

where Mi

The mass equations which

of propellant

as a function

values

mass

procedure loading

hydrogen

18 and

In both

cases,

optimum

inert

production

fraction

propellants

first

stage stage

engines

predicts 215

without

shown locally is

done

are

on

a fifth-order

polynomial

versus

nitrogen

mass

calculations,

from LEO

fraction

was fit to the

shown

in figure

4.

calculated

by

to account

for the inert

a similar

is then just the sum of the

This

sec.,

the specific

for engine mass

The actual

with

impulse

impulse

optimum

inert

of inert

up to an

estimated

the

higher,

for

added.

using

for the 3400

of either

is slightly

Earth,

injection,

the result

specific

m/sec

hydrogen

or

as expected,

and tank mass. only,

assuming

If both fuel and oxidizer these

from

of the amount with

is consistent propellant

below

to be brought

decreases

that the optimum

Mars,

modified

as a function

required

to 70%.

hydrogen

was

of hydrogen

respectively,

well

addition

injection

these

90%.

propellant

amount

inert injection.

on Mars.

the inert

from

stages.

of hydrogen

of 60%

can be calculated

To calculate

to be brought

the required

(3) does not account

values

manufactured

impulse

for the second

amount

impulse,

engine

from 0% to approximately

in the

fueled

AV is about

since equation The

the

(3), which

stage

specific

19 show

impulse.

specific

for the f'n'st and second

and methane

equation

specific

The total mass of hydrogen

needed

Figures

methane

of hydrogen

level.

for the first stage

propellant

inert levels

to that used

amounts

necessary

of propellant

This was done using The

AV ]

of propellant

will give a specified

calculated

first

exd

masses 11

are

that

the rest

are brought higher

of the propellant

from Earth, by

a

factor

is

and only inert of

7

(for

i*

The

H2-O2) and 17.6 (for CH4-O2).

location

and relative

value

of the minimum

does

not

change.

2015 The

"2015

low Martian was

Mission"

orbit (LMO)

600 kg, with

assumed.

scenario

a thrust

This results

payload

propellant

For

level

Isp

H2-O 2 propellant

addition

of 50%

nitrogen,

missions

were

fraction 02

values

mass

and

of 1,800

for propellant This

missions,

Table

assumed

Three

kg and

engine

seconds

waiting

the engine

engines

a total

with

for the

reduced

impulse

analysis

per

thrust

Table

in

mass

mission

were

of 400 kN.

IX shows the

injection

CH4-O 2 propellant.

and eqns.

X shows

no inert

to 363 sec. and

was performed

mass.

while

for this

were

specific

it was

vehicle

The

is 5,320 m/s.

404

values

is 11% of the propellant

propellant

kN per engine.

by CEC

these

calculated.

propellant

5-7, mass

306

breakdowns

the assumption

the mass

breakdown

breakdown

495

With

the

sec. respectively.

using

mass

was

for these

that the tank for the two H 2-

for

the

two

CH4-O 2

missions.

Again,

accelerations

the feasibility using

engine

predicted

for the

these

of 133.5

launch to a return

of this study,

kg, and the AV required

seconds

Using

a single-stage

the purpose

in a total

mass is 5,400

The

1°.

describes

MISSION

the

assumed

Figures

for the beginning

in regards

to crew

thrust

level

of 400

and end of the engine

and equipment kN

with

the

safety. derived

This system

bum to insure was performed masses.

The

in table XI.

of hydrogen

fueled

function

total

are shown

mass

methane-

calculated

of the scenario

accelerations The

were

"2015"

which

mission

20 and 21 show

of the percentage

must

be brought

scenarios

the amount of inert

from

were calculated of hydrogen

loading

level.

LEO

for the

hydrogen-

by the same method

which

must

be brought

For this mission,

there

and

the

used earlier. from

LEO

as a

is no advantage

to

inert injection. This is slightly optimum engine

Isp and

decrease While

surprising,

for this case tank

mass

the required there

since should

is not result

the simple be about

estimation

340 sec.

in an accurate

using The

equation

difference

estimate

of whether

(3) suggests shows

that the

that neglecting

inert

injection

will

hydrogen.

is no performance

advantage

of inert injection,

12

the fact that inert

addition

has

4

little effect This

on performance

indicates

that

manufacturing

high

processes

inert

mission than

injection

as well.

propellant

purity

fraction

is not

would

have

However,

propellant

also resulted

this would

While

type is not currently

there

under

of inert is itself

a critical

was done only at the beginning

the one assumed.

engine

50% mass

for Mars-manufactured

If the inert injection low,

up to nearly

is no reason

parameter

in the

from a performance of the flight,

in a decreased

have required

a significant

when

hydrogen

the assumption

result.

design

standpoint. the required

requirement

Isp

could

is

for this

of a more complex

why such an engine

of

engine

not be made,

this

study.

CONCLUSIONS A novel

Mars mission

atmosphere return

requirement to analyze

This

for fuel brought

from Earth.

A standard

of inert injection

from the Mars surface

missions

were only representative

staging

criterion

specifically injection only

eliminate The would

when would

to take

can be varied

analyses

during

clearly

the amount

beginning

be interesting

engine

analyzed.

strategies

increase

the Martian

chamber

of an Earth-

but can result combustion

in a lower

code

For one ease,

was found

to result

boost

of higher

This could

allow

was run

a two-stage in a reduced

considered. with

Design

low Av; choosing

leverage

For example,

using

utilizing

advantages

but

The advantage

of new mission

in propellant

the

injection,

of the

the

strategies inert

gas

inert injection injection

but

mass carded.

using

assuming result

being

on stages

advantage.

would

all calculated

to do the calculations the boost.

is used

by this strategy.

propellant

use of inert propellant

currently

of the increase

gained

of the

were

optimum

inert injection

advantage

most of the penalty mission

were

the inert injection

to make

cases of mission

carefully

may increase the

not designed

is greatest

intended

during

rocket

used,

from

performance.

combinations to Earth,

inert gasses

from Earth. were

of inert injection

of propellant

on engine

and two propellant

liquefied

into the combustion

mass

of fuel brought These

to inject

in a greater

the effect

direct

mass

where

results

Two missions

mass

was studied,

are used as inert reaction

vehicle.

launch

strategy

a fixed

of inert

that the inert injection,

in additional

13

fraction

mass

savings.

in each and hence

stage.

It

the Isp,

REFERENCES 1.

French,

J.R.,

Interplanetary 2.

Stancati, Report

3.

o

5.

7.

R.M.,

August,

1990.

Baker,

Vol. 42, pp.

167-170,

Niehoff,

J.C.,

and

Landis,

G.A.

and

D.,

October,

1988.

Leitman,

G.,

Zuppero,

R.M.,

"Mars

147-160,

1989.

D.L.,

"Acetylen

Transfer,"

and

Chap.

Landis,

1st. Annual AZ.

Gordon,

S.,

Chemical

Equilibrium

McBride,

December

to Mars

Direct,"

from

1988,"

NASA

of Rockets," Japan,

1961;

Final

America,

Interplanetary

Atmospheric

CO 2 on

Mars,"

1992.

Technical

Memorandum

International

in expanded

G.A.,

B.J.,

"Optimum

Rocket

form

on Resources

"Computer

Propulsion

100470,

Symposium

on

in G. Leitman

Program Performance,

Detonations,"

NASA

SP-273,

"Orientation/lst

Intern

1989.

14

for

of Near-Earth

Rocket

Co.,

in Aerospace

of the British

3rd.

System

1978.

in 1999!,"

Journal

e Fuel

Refueling

October,

Compositions,

B.; Boeing

of the British

(ed.)

13, pp. 377-388.

Symposium

and Chapman-Jouguet

37857,

Tokyo,

Techniques,

Tucson,

i0. Donahue,

Efficiency

Journal

Automated

29 No. 2, pp. 294-296,

of Mars,

"Propulsive

and

"Humans

and Rockets,Vol.

"Environment

"Mars

No. SAI 1-120-196-$7,

Zubrin,

Linne,

Resources,"

1989.

W.C.,

D. A.,

and Astronautics,

A.,

Martian

Baker,

of Spacecraft

Kaplan,

and Wells,

Report

Vol. 42, No. 4, pp.

Optimization

9.

Society,

and

D.A.,

Rockets

8.

from

and Presentation,"

Journal 6

Propellants

M.L.,

Zubrin,

Soc.,

"Rocket

for

Space,

Briefing,"

7-10

Calculation

Incident

March,

Energy-Limited July,

1991,

of Complex

and Reflected

Shocks,

1976. NASA

Contract

NAS8-

d

Table I. Compositionof theMartianAtmosphere (from ref. 6)

Table

IV

CH4-O

2 Second

Stage

Propellant

Gas CO2 N2 Ar 02 (X) Ne Kr Xe

Table

Percent(by mass) 96.8 1.7 1.4 0.1 0.04 1.1ppm 0.6ppm 0.2 ppm

Tank Engine Capsule

Example

Propellant

Table V Second Stage

Flow-rate

0.2143

kg/sec

:

inert in oxidizer

0.0357

kg/sec

:

inert in fuel

0.6428

kg/sec

:

oxidizer

0.1072

kg/sec

:

fuel

1.0000

kg/sec

:

Total

Table

III Second

Stage

Propellant Tank Engine Capsule

:

TOTAL

Breakdown

13,718

kg

:

181 kg

:

12,250

kg

35,294

kg

Accelerations Final Acceleration

H2-O2

0.86gMars

1.72gMars

0.33gEarth

0.66gEarth

0.67gMars

1.64gMars

0.26gEarth

0.63gEarth

Table

VI

H2-O2

First

kg

1,372 kg

:

181 kg

:

2,078

Propellant

:

and payload:

:

Acceleration

(I-I2)

Mass

kg

Propellant

CH4-O2

H2-O2

20,784

Initial

(O2)

Breakdown

:

and payload: TOTAL

II

Mass

Stage

Mass

Breakdown

0 % Inert Propellant

75 % Iner_

34,135

kg

110,838

kg

3,413

kg

11,084

kg

2,722

kg

2,722

kg

Payload

27,521

kg

27,521

kg

TOTAL

67,791

kg

152,164

kg

12,250

kg

Tank and Structure

27,521

kg

Engine

Isp=495 sec (no inert); 266 see (75% inert).

15

Table

Table

VII

_4.4-O

2 First

(AV = 3,400

Stage

Mass

Breakdown

m/s) 0% Inert

Propellant

IX

H2-O2

Propellant

(single

stage to Mars orbit) 0% Inert

75% Inert kg

Propellant

1,969

kg

kg

2,722

33,348

kg

98,279

kg

56,572

kg

Tank and Structure

5,637

kg

Engine

2,722

Payload TOTAL

Mission

Mass

Breakdown

50% Inert

18,405

kg

40,430

kg

Tank and Structure

2,024

kg

4,447

kg

kg

Engine

1,800

kg

1,800 kg

33,348

kg

Payload

5,406

kg

5,406

kg

252_660

kg

TOTAL

27,635

kg

52,083

kg

196,900

Isp--404 sec (no inert); 229 set: (75% inert) Table

X

CH4-O2 Table

(single

VHI

First Stage

Accelerations

(AV= 3,400

Propellant

[ H2-O2

gMars

0% Inert

m/s)

gEarth

with 0% Inert

A (initial)

i9.68

5.26

2.01.

A (final)

39.65

10.60

4.04

with 75% Inert A (initial)

8.77

2.34

0.89

A (final)

32.29

8.63

3.29

CH4-O2

with 0%Incr_

A (initial)

13.58

3.63

1.38

A (final)

31.98

8.55

3.26

with 75% Inch A (initial) A (final)

5.28

1.41

0.54

23.93

6.40

2.44

16

Mass

Breakdown

stage to Mars orbit)

Propellant m/s 2

Mission

50% Inert

29,553

kg

76,603

kg

Tank and Structure

3,251

kg

8,426

kg

Engine

1,800 kg

1,800

kg

Payload

5,406

kg

5,406

kg

TOTAL

40,010

kg

92,236

kg

Table XI 2015MissionAccelerations H2-O2

wiqh 0%

gMars Inea

gEarth

A (initial)

14.49

3.87

1.48

A (final)

43.37

11.60

4.42

with 50% Inert 7.69

2.06

0.78

A (final)

34.35

9.19

3.50

CH4-O2

with 0%Inert

A (initial)

A (initial)

10.01

2.68

1.02

A (final)

38.28

10.24

3.90

with 50% Inert A (initial) A (final)

4.34

1.16

0.44

25.61

6.85

2.61

17

500

INERT Ok) @3

4OO

A

N2

[]

Ar

O CO2

m

E

tam

¢,J

300

J 200 C_

Pc = 200

psia

100 0

20

40

60

80

100

% Inert in Propellant

Figure 1. Effect on propellant specific impulse of addition of inert reaction mass to H2-O2 fueled engine

425

A

¢0

N2

OAr CO g_

[]

325

CO2

E ¢J .m ¢,0

E

225

g_ ¢,0

Pc = 200 psia 125 0

20

40 % Inert

60 in Propellant

80

100

Figure 2. Effect on propellant specific impulse of addition of inert reaction mass to CH4-02 fueled engine

18

350

I

'

I

I

I

INERT n

N2

,, Ar C02

300J

250

S :zoo

g Pc = 200 psia I

150

|

I

20

0

|

40 % Inert

I

1

1

60 in Propellant

00

80

Figure 3. Effect on propellant specific impulse of addition of inert reaction mass to CO-O2 fueled engine

500

VUEL --

H2

........... .....

CH4 CO

% %

% % %

% %

Inert

- N2

% %

100 0

20

40 % Inert

60 in Propellant

80

100

Figure 4. Comparison of effect on propellant specific impulse due to the addition of local inert to hydrogn, methane, and carbon monoxide fueled engines

19

3000

2500 H2

glO

.............. .....

_- 2000 E

"_

1500

_

1000

E o Z

500

N

CH4 CO L

Inert

- N2

0 0

20

40

60

% Inertin

Propellant

80

100

Figure 5. Isp(F+O) [Specific impulse normalized to fuel and oxidizer flow], showing effect of addition of inert reaction mass to hydrogen, methane, and carbon monoxide fueled engines

5OO H2-O2

Propellant

"-" 400 J_ e_

E _

o_

E =

300

200

¢d m

--

5.0

......... .........

6.0 3.5

....... .........

7.0 8.5 Inert

- N2

100 0

20

40 % Inert

60 in Propellant

Figure 6. The effects on propellant

80

100

specific impulse of addition of N2 to H2-O2 fueled engine while varying the O/F mixture ratio ( Exit Area Ratio Constant at 200 ) 20

500

|

!

CH4-O2

g

I

Propellant

!dJc L.Q ""

3.4

400 ........

r_

4.5

sJ =

1.7

300

S =

21111 Inert

- N2

%"_%__

100 0

20

40 % Inert

60 in Propellant

80

I00

Figure 7. The Effects on propellant specific impulse of addition of N2 to CI-I4-O2 fueled engine while varying the O/F mixture ratio ( Exit Area Ratio Constant at 200 )

350 CO-O2

Mixture

O/F

'--

.550

............ ......... ....... .........

£75 .775 .425 .325

Propellant

Inert

- N2

I00 0

20

40 60 % Inert in Propellant

80

100

Figure 8. The effect on propellant specific impulse of addition of N2 to CO-O2 fueled engine while varying the O/F mixture ratio ( Exit Area Ratio Constant at 200 ) 21

3OOO H2-O2

Propellant

2500

di,xalxa__0 i

w_

5.0

g_

E

2000

............. ......... ....... .........

)mr ¢d

1500

tLI

N

.i i ¢I

6.0 3.5 7.0 8.5

1000

E L.

Zo

500 Inert

- N2

0 0

20

40

60

80

100

% Inert in Propellant

Figure 9. Isp(F+O) [Specific impulse normalized to fuel and oxidizer flow], showing effect of addition of inert reaction mass to hydrogen fueled engine while varying the mixture ratio ( Exit Area Ratio Constant at 200 )

3OOO

I

"

i

"_

I

1

CH4-O2

Mixture

O/F

2500 2000

............

4.5

1500

......... ....... .........

2.5 5.3 3.4 1.7

I

Propellant

t

}

"gV

:_ Z

1000

soo Inert ,

0 0

I

20

,

I

,_

40 % Inert

I

60

_

- N2 '

80

•

100

in Propellant

Figure 10. lsp(F+O) [Specific impulse normalized to fuel and oxidizer flow], showing effect of addition of inert reaction mass to methane fueled engine while varying the mixture ratio ( Exit Area Ratio Constant at 200 ) 22

2500 CO-O2

2000

KLxIlgg..fl

1500

.......... .........

.550 .675 .775

....... .........

.425 .325

gh

r.j om

¢la gg

Propellant

1000 6.1 .am ,la

E [_. o

500 Inert

- N2

0 0

20

40 60 % Inert in Propellant

80

100

Figure 11. Isp(F+O) [Specific impulse normalized to fuel and oxidizer flow], showing, effect of addition of inert reaction mass to a carbon-monoxide fueled engine while varying the nuxture ratio ( Exit/u'ca Ratio Constant at 200 )

600

,

, H2-O2

400

,

""

300 _=

_

500

.........

400

200I- ....... 100

, Propellant

/t -l

0

200

"....." ".

100 ,"

20

,

Inert |

,

|

40 60 % Inert in Propellant

- N2 ,

|

80

,

100

Figure 12. The Effects on propellant specific impulse due to the addition of nitrogen to a hydrogen-fueled engine While Varying the Exit Area Ratio ( Mixture Ratio - 6.0 ) 23

500

"-"

CH4- 02 Propellant

400

E _= 300

Area

e_

Ratio

e__

500 ............ ......... ....... .........

E = 200 t.t

400 300 200 100 Inert

- N2

100 0

20

40 60 % Inert in Propellant

80

100

Hgurc 13. The Effects on propellant due to the Addition of motrpgem While Varying the Exit Area Ratio ( Mixture Ratio - 3.4 )

400

--

•

I

I

22.'..'.2....

CO-O2

I

to a methane fueled engine

l

Propellant

300 ._

d 500 E= 200

........... .........

400 300

:::"::"

_

,

100 0

I

20

""_ "_

Inert-N2 ,

I

40 % inert

,

I

60 in Propellant

.

_ I

80

,

100

Hgure 14 The Effects on Isp Due to the Addition of N2 to a carbon monoxide While Varying the Exit Area Ratio ( Mixture Ratio - 0.55 ) 24

fueled engine

3000 H2-O2

Propellant

"0

m

E

Area 2000 ......... .........

500 400 300

....... .........

200 100

Imm

_N

1000

g_

m

E 0

z

Inert

- N2

0 0

Figure

20

40 60 % Inert in Propellant

80

100

15. Isp(F+O) [Specific impulse normalized to fuel and oxidizer flow], showing effect of addition of inert reaction mass to a hydrogen fueled engine While Varying the Exit Area Ratio ( Mixture Ratio - 6.0 )

25

3OO0 CH4-O2

Propellant

g 2000

e 500 _'_

400

......... .......

300 200

j" . jP"

/-

1000

Z

Inert I

0

E

I

40 % Inert

I

,_,J

•

•

60 in Propellant

"

80

100

[Specific impulse normalized to fuel and oxidizer flow], showing effect of addition of inert reaction mass to a methane fueled engine ( Mixture Ratio - 3.4 )

30OO

O.

I

20

0

Figure 16. Isp(F+O)

I

- N2

'

CO'O- 2 Propellant '

'

2000

lain

500

¢o

_o

el_

1000

..........

400

.........

300

['-..7"

200

J _/ /

t_

E i@

Z

Inert I

0 0

Figure

I

20

I

,

I

40 % Inert

•

I

60 in Propellant

•

- N2 I

80

•

100

17. Isp(F+O) [Specific impulse normalized to fuel and oxidizer flow], showing effect of addition of inert reaction mass to a carbon-monoxide fueled engine While Varying the Exit Area Ratio (Mixture Ratio - 0.55 ) 26

7000

!

_

I

•

!

I

H2-O2-N2 Propellant Zubrin

"w

*_

6800

.a E

6600

o

6400

1

6200 o k

,_ = zoo om = 6.0

6000

Enllne Thrust

i :

1,354

,

5800 0

\

_

2.,721 tl kN

I

_

=

20

/

I

i

40

I

60

1

t

,

I

80

100

% Inert

Figure

18. Hydrogen Mass required to Support the "Direct" return launch (two-stage surface to Earth Injection) Using H2-O2-N2 Propellant Mixture

launch from

4300 CH4-O2-N2 Propellant Zubrin ¢_ 4200

_ E

4100

o

& =

4000

3900 Pc = 20.68 AR : 200 O/l _ = 3.4

"o

¢

3800

Engine Thrust

MPa

Mm = 2,722 : 1,354 kN

kg

3700 0

20

40

60

80

100

% Inert

Figure 19. Hydrogen Mass required to Support the "Direct" return launch (two-stage launch from surface to Earth Injection) Using CH4-O2-N2 Propellant Mixture 27

40000

!

I

H2-O2.N2

Propellant 2015

I ,4

_e v

30000

I

E O

&

,=1

20000

t_ Pc = 20.68 AR : 200 Off' : 6.0

10000 'O

MPi

Engine Thrust

Mass = 400

i

I

I

- l,g00 kN

1

kg

I 0

0

i,,

I

20

i

40

i

•

t

60

80

I00

% Inert

Figure 20. Hydrogen

Mass required to Support the "2015 mission" return launch scenario mars orbit) with H2-O2-N2 Propellant Mixture

200O0

|

(surface to

I

CH4-O2-N2 Propellant 2015 CaD dg

15000 E o W_

10000

¢w Pc : 20.68 AR = 200 O/F = 3.4 Engine Thrust

5000

0

,

0

Figure 21. Hydrogen

MPi

Mass . 400

= 1,800 kN

kg

I

,

20

,

I

40 % Inert

i

I

60

,

80

Mass required to Support the "2015 mission" return launch (surface to mars orbit) with CH4-O2-N2 Propellant Mixture 28

REPORT

DOCUMENTATION

Form Approved OMB NO. 0704-0188

PAGE

' Public repotting burdan for this collection oi information iS estimated to average'1 hour per response, including the time for reviewing instructions, searchLng existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing (his f_._rden. ¢o Washington Headquartars Services. Directorat_ for information Operations ant/Reports, t2f5 Jefferson Davis Highway, Suita 1204, Arlington, VA 22202-4302, and to the Office of Management and Bu0get, Paperwork Reduction Project (0704-0188). Washington. DC 20503,

I 1. AGENCY USE ONLY (Leave blank)

2. REPORT DATE

3. REPORT TYPE AND DATES COVERED

September

1992

Final

4. TITLE AND SUBTITLE

Contractor

Report

5. FUNDING NUMBERS

Effect of Inert Propellant Injection Mars Ascent Vehicle Performance

on %

WU-506-.41-11 ¢

AUTHOR(S) James

E. Colvin

and Geoffrey

A. Landis

7. PERFORMING ORGANIZATION NAME(S) AND ,_DDRESS(ES) Sverdrup Lewis

Technology,

2001

Research

Brook

Park,

Inc.

Center

Aerospace

Group E-7312

Parkway

Ohio

44142

9. SPONSORING/MONITORING

Ohio

10. SPONSORING/MONITORING AGENCY REPORT NUMBER

AGENCY NAMES(S) AND ADDRESS(ES)

National Aeronautics and Space Lewis Research Center Cleveland,

8. PERFORMIN(_ ORGANIZATION REPORTNUMBER

Administration NASA

CR-189238

AIAA-92-3447

44135-3191

11. SUPPLEMENTARY NOTES Prepared for the 28th Joint Propulsion Conference and Exhibit cosponsored by the AIAA, SAE, ASME, and ASEE, Nashville, Tennessee, July 6 - 8, !992. James E. Cotvin, University of Arizona Space Engineering Research Center, Tucson, Arizona 85712. Geoffrey A. Landis, Sverdrup Technology, Inc., Lewis Research Center Group, 2001 Aerospace Parkway, Brook Park, Ohio 44142. Responsible person, Geoffrey A. Landis, (216) 433-2238. 121. DISTRIBUTION/AVAILABILITY

Unclassified Subject

13.

12b. DISTRIBUTION CODE

- Unlimited

Category

ABSTRACT

(Maximum

A Mars

ascent

the vehicle in-situ

STATEMENT

vehicle

resource

is limited can

and does atmosphere

step. The

combustion

words)

performance

of the Martian processing

200

code

expansion

ratio,

candidate

missions

decreased

using

effect

not have which

oxidizer/fuel to determine propellant

to be brought

gas injection

Earth.

on rocket

injection

the required

gas

which into

can be brought

the engine,

Carbon

dioxide,

by compressing

equilibrium

and inert

how

inert

from

be separated

chemical ratio,

by the propellant

by injecting

could

of inert

that calculated

inert

in performance

be improved

engine

performance of varying

fraction.

mass

Results

of return

nitrogen,

the atmosphere,

for engines

was

and argon

analyzed

combustion

needed

were

in low

engines;

In some

cases

as an

are constituents

any chemical with

a numerical

chamber

pressure,

applied Earth

to several

orbit

could

be

injection.

15. NUMBER OF PAGES

14. SUBJECT TERMS Rocket

Earth.

gas is available

without

of this analysis

propellant

from

if the inert

Mars

missions;

Space

3O

exploration

16. PRICE CODE

A03 17. SECURITY CLASSlRCATION OF REPORT Unclassified NSN 7540-01-280-5500

18. 'SECURITY CLASSIFICATION OF THIS PAGE Unclassified

19. SECURITYCLASSIFICATION OF ABSTRACT

20. LIMITATION OF ABSTRACT

Unclassified Standard Form 298 (Rev. 2-89) Prescribed 298-102

by ANSI

Std, Z39-1

B