ESTIMATION OF THE
DYNAMIC FLUID PRESSURE EXERTED ON THE LUNAR SURFACE BY THE DESCENT
ENGINE EXHAUST OF THE APOLLO 12 (LM6) LUNAR MODULE AT HOVER THRUST
IMMEDIATELY PRIOR TO LANDING
This page derives an estimate of the dynamic fluid pressure upon
the lunar surface from the impingement of the descent propulsion
system (DPS) exhaust plume of the Apollo 12 lunar module (LM) as it
descended the last few feet to touchdown.
The DPS engine could be throttled between 10% and 94% of its
thrust capacity in order to achieve continuous variable descent rates
consistent with the nominal descent profile. At 94% thrust, the DPS
engine produced 9,870 lbf (43,902 N)
thrust. [ANR, p. 92] This thrust level was used for prelanding
orbital maneuvers and was available for various abort modes. It is
not the thrust setting used in the final phase of a normal touchdown.
In order to hover, ascend at a constant rate, or descend at a
constant rate, the force of engine thrust must balance the force of
gravity upon the spacecraft, i.e., the spacecraft's weight in lunar
gravity. More thrust than weight would produce an ascent at an
increasing rate.
1. ESTIMATION OF
LM6 LANDED WEIGHT IN LUNAR GRAVITY
Approximately half the LM's mass at launch from earth was DPS
propellant in the form of Aerozine 50 and nitrogen tetroxide contained
in redundant tanks in the LM descent stage. This propellant was
consumed and almost fully depleted by the DPS during the powered
descent. At touchdown only a small fraction of this mass remained.
Other consumables were partially depleted prior to touchdown.
The table below gives estimates of the major LM components and
consumables at touchdown. Data comes from the AS12 loadout
[Reports12, p. 44] and from the consumables tracking table [Reports12,
p. 137].
LM6 LANDING MASS

ITEM 
MASS lbm

MASS kg

Ascent stage dry^{1} 
4,760 
2,164 
Ascent propellant 
5,170 
2,350 
Descent stage dry^{1} 
4,832^{2} 
2,196 
Descent propellant 
705 
320 
RCS propellant 
350^{3} 
159 
Crew 
293 
133 
TOTAL 
16,110 
7,322 
^{1} Dry masses include spacecraft structure,
equipment, water, oxygen. They exclude propellant and crew.
^{2} Differs from published amount due to estimation of
water depletion for cooling prior to landing. Mission report
specifies 171.2 lbm (77.8 kg) descent stage water was consumed
prior to LM ascent stage liftoff for return to earth. We
estimate onefourth of this, 42.8 lbm (19.5 kg), may have been
used during the descent.
^{3} Published figure is 150 lbm (68.2 kg) at LM ascent
stage impact prior to returning to orbit, out of an initial load
of 549 lbm (250 kg). We estimate half of RCS propellant
consumed during mission was consumed prior to touchdown.

Earth gravity exerts a force of 1 lbf for each pound of mass, (9.8
N for each kilogram). Lunar gravity exerts 0.16 times that force. We
compute the force of lunar gravity on the LM at touchdown as 2,690 lbf
(11,900 N).
2. ESTIMATION OF
HOVER THRUST AND ENGINE OPERATING PARAMETERS
A landing weight of 2,690 lbf (11,900 N) dictates an equivalent
hover thrust of 2,690 lbf (11,900 N). 1,050 lbf (4,670 N) is
considered 10% thrust, while 9,870 lbf (43,902 N) is 94% [ANR, p. 92],
indicating a maximum rated thrust of 10,500 lbf (46,704 N) for the
DPS. 2,690 lbf (11,965 N) corresponds to 25.6% rated thrust. The
hover thrust is specified as "approximately 25%" rated thrust [ANR,
p. 102].
Exhaust exit velocity (V_{e}) is related to thrust (F) by
F = m° V_{e}
where m° (pronounced "mdot") is the mass flow rate of
propellants to the engine in units of mass per unit of time. The mass
flow rate of most rocket engines is a constant because the thrust is
fixed. The DPS is a variablethrust engine. Its thrust is controlled
by varying its mass flow rate.
Unfortunately we cannot locate any published values for the DPS
m° at any throttle setting. We can, however, deduce it from other
measures provided. The description of the DPS propellant quantity
gaging system reads in part
"The lowlevel sensors provide a discrete signal to cause
the warning light to go on when the propellant level in any tank is
down to 9.4 inches [239 mm] (equivalent to 5.6% propellant
remaining). When this warning light goes on the quantity of
propellant remaining in sufficient for only 2 minutes of engine burn
at hover thrust (approximately 25%)." [ANR, p. 102]
This allows us to equate a known mass of fuel with a known operation
time at the throttle setting we are most interested in. The DPS fuel
quantity is given in the AS12 loadout as 17,925 lbm (8,148 kg)
[Reports12, p. 32] and as 17,489 lbm (7,950 kg) in the AS12
consumables report [Reports12, p. 133]. The latter quantity is
qualified as "usable" propellant, indicating that the difference of
436 lbs (198 kg) represents fuel which was loaded but which cannot be
delivered to the engine due to limitations of the propellant feed
equipment ("tankage"). 5.6% of the usable fuel mass is 979 lbm (445
kg). At approximately 25% throttle, this mass would be used in 120
seconds, suggesting a mass flow rate m° = 8.16 lbm/sec (3.71
kg/s) at 25% throttle. Because this is a computed estimate based on
informally reported information, we must take steps to verify it.
The ascent propulsion system (APS) and the reaction control system
(RCS) jets use the same fuel combination as the DPS and their mass
flow rates are published along with the thrust they produce. APS
m° is 11.2 lbm/sec (5.1 kg/s). Its nominal thrust is given as
3,500 lbf (15,568 N) [ANR, p. 93]. The specific impulse
(I_{sp}) of a particular propellant is the ratio of thrust to
mass flow rate, giving an indication of how propulsive a particular
propellant formulation is.
I_{sp} = F / m°
For the APS I_{sp} = 312 sec. The RCS m° is 0.357 lbm/sec
(0.162 kg/s) and the thrust is 100 lbf (445 N). I_{sp} in
this system is 280 sec. Values of 285370 seconds are expected for
this type of propellant. At 25% thrust we estimate a DPS m°
of 8.16 lbm/sec (3.71 kg/s) and a force of 2,690 lbf (11,965 N) for
an I_{sp} of 330, within the accepted range.
We can therefore be confident that the mass flow rate of the DPS
at 25% thrust, 8.16 lbm/sec (3.71 kg/s), is a reasonable value.
3. COMPUTATION OF
EXHAUST VELOCITY
As previously stated, exhaust velocity (V_{e}) is related to
mass flow rate (m°) and thrust (F) by
F = m° V_{e}
86,618 lbf = (8.16 lbs/sec) V_{e}
11,965 N = (3.71 kg/s) V_{e}
V_{e} is 10,615 fps (3,225 m/sec).
As a doublecheck, we note that V_{e} = I_{sp}
g_{0}. (32.1 ft/sec^{2}) (330 sec) = 10,593 fps, or
(9.81 m/s^{2}) (330 sec) = 3,237 m/s. Clearly all the
relevant rocketry equations closely agree with our estimate.
4. COMPUTATION OF
DYNAMIC FLUID PRESSURE
The static pressure of a gaseous fluid is determined by the
kineticmolecular theory and is computed for a given temperature,
volume, and molar mass.
The dynamic pressure of a fluid (gaseous or liquid) in the
plumeimpingement case is the force exerted on a flat surface by the
motion of a fluid, not its kineticmolecular state. It is
given by
F = m° V_{e}
which is the same as the thrust equation above. Thus the values
computed in section 2 can be used here: 2,690 lbf (11,965 N).
To convert this to unit pressure, assuming an ideal plume, we
divide total dynamic fluid pressure by the area of impingement equal
to the area of the exit plane of the DPS nozzle, 19.0 square feet
(1.77 square meters). This value is 141 lbf/ft^{2} (6,760
Pa), or approximately 0.98 psi.
5. ESTIMATION OF
THE EFFECT OF PLUME DISPERSAL IN A VACUUM
As stated above, the actual plume differs from the theoretical
plume used to derive average plume mass density in that it disperses
in a vacuum. Dispersal has two primary effects: a decrease in exhaust
velocity and a decrease in density. Dispersal is a function of
distance from the exit plane.
Molecules exit the nozzle on similar but not identical
trajectories. The flow is essentially collinear at the nozzle exit,
but collisions between molecules will soon eliminate coherence in the
flow. While the plume conserves kinetic energy, that energy is no
longer directed and the dynamic fluid pressure law does not apply.
The departure of gas molecules from their collinear trajectories
causes a decrease in the density of the gas. The less collinear the
flow, the greater the probability a wayward molecule will strike a
neighbor and change its trajectory. However, the more molecules that
depart from collinear trajectories, the more likely they will adopt
divergent trajectories, lowering the probability of subsequent
collisions.
The fluid velocity as applicable to dynamic pressure will
eventually fall to zero at some distance from the nozzle plane, as all
molecular activity becomes random and static pressure laws take
precedence. The fluid density will eventually fall to zero at some
distance from the nozzle plane as molecules on divergent trajectories
depart the plume volume. Thus the theoretical fluid pressure at the
exit plane is a maximum, and an attenuation factor can be formulated
as a function of distance (h) from the exit plane.
When h = 0 the attenuation function a(h) must be 1.0. Commander
Charles "Pete" Conrad reported that the lunar surface dust began to be
visibly displaced when the LM had descended to an altitude of 300 feet
(91 m) [Reports12, p. 169]. Thus the derived attenuation function
a(h) must allow for significant dynamic fluid pressure at h = 300 ft
(91 m). We arbitrarily set this significance threshold at 1 pound per
square foot (48 Pa) at h = 300 ft (91 m).
We have already theorized that the majority of the dispersal will
occur shortly after exit. Therefore a(h) must fall off rapidly
initially, and more slowly as h increases.
