3.5 A-4 STAGE ENGINE

For the A-4 stage, two engines of $7500-$ pound thrust each were selected, for a combined thrust of 15000 pounds. It is assumed that the mission assigned to this fourth and last stage of

Figure 3-7.-48K A-3 stage propulsion system preliminary design layout.

our space vehicle may require prolonged cruising periods prior to ignition and possibly even longer waiting periods prior to reignition. While it would be desirable to utilize the high-energy propellants of the second and third stages, the fact that they are cryogenics poses some prob- lems. Although cryogenic propellants could probably be used with refined insulation techniques, they were not selected because of the systems complication for a vehicle of this size. Solid propellants were also ruled out because of the need for repeated starts and throttling.

Figure 3-8.-A-3 stage engine and propulsion system operational sequence.

A hypergolic, storable propellant combination possesses certain characteristics which contribute to high reliability. Among these are simplicity of ignition and ease of propellant maintenance, since the propellants can be contained in closed vessels over reasonable temperature ranges for considerable periods of time without developing excessively high pressures, or undergoing unacceptable changes in composition.

Among the applicable storable propellant combinations with high performance are chlorine trifluoride $\left(\mathrm{ClF}_{3}\right) /$ hydrazine $\left(\mathrm{N}_{2} \mathrm{H}_{4}\right)$ and nitrogen tetroxide $\left(\mathrm{N}_{2} \mathrm{O}_{4}\right) /$ hydrazine. Hydrazine, as a monopropellant, is prone to explosive thermal decomposition. However, the condition can be remedied by certain additives. The $\mathrm{ClF}_{3} / \mathrm{N}_{2} \mathrm{H}_{4}$ combination has slightly higher performance than the $\mathrm{N}_{2} \mathrm{O}_{4} / \mathrm{N}_{2} \mathrm{H}_{4}$ combination. Handling of $\mathrm{ClF}_{3}$, however, requires special design provisions because of its thermal characteristics. For this reason, the $\mathrm{N}_{2} \mathrm{O}_{4} / \mathrm{N}_{2} \mathrm{H}_{4}$ propellant combination was chosen for the A-4 engine. It is worthy of note that the performance of $\mathrm{N}_{2} \mathrm{O}_{4} / \mathrm{N}_{2} \mathrm{H}_{4}$ is comparable to that of $\mathrm{LO}_{2} / \mathrm{RP}-1$.

Teflon and Teflon 100X can be used as seal material in the A-4 engine system. Kel-F, while a satisfactory material for use with $\mathrm{N}_{2} \mathrm{H}_{4}$, degrades after short-term service in $\mathrm{N}_{2} \mathrm{O}_{4}$. Most series 300 stainless steels, aluminum allovs, nickel, and nickel-base brazing alloys can be used as construction materials.

General Engine System Description

The A-4 engine is a multiple-start, variablethrust, gimbaled, bipropellant system. The basic system includes a thrust chamber assembly utilizing combined ablative and radiation cooling, propellant ducts, valves, and control subsystems. Thrust chamber ignition is achieved by the hypergolicity of the propellants. One significant feature of this engine system is the clustering of two thrust chambers to one propellant feed system and one set of propellant controls. The propellants are fed by pressurants directly from the propellant tanks through the main propellant valves to the thrust chamber inlets. Gaseous helium supplied from high-pressure bottles is used for pressurization of both tanks. The pressurant is heated in heat exchangers located at the thrust chamber nozzle extensions before expansion through a pressure regulator and transfer to the propellant tanks. Helium gas is also used to operate the main valves and the gimbal actuators.

A-4 engine operating parameters at vacuum condition are presented in table 3-5. The engine and propulsion system schematic diagram is shown in figure 3-9.

The engine gimbal blocks are fastened to thrust mounts which are attached to the aft end of the oxidizer tank. The fuel tank is attached forward of the oxidizer tank to form an integral vehicle structure. As in the A-3 system, the thrust loads are transmitted to the payload through the pressure-stabilized tank assembly. The propellant ducts between fuel tank and engine systems are routed outboard and covered by fairings for protection against aerodynamic heating and for lower air resistance during first-stage boost.

Both throttling and propellant-utilization control are achieved by varying the degree of opening of both propellant valves. The positions of the valves are controlled by the vehicle guidance system in conjunction with a vehicle propellant quantity measuring system. Thrust vector control is accomplished by gimbaling the thrust chambers. The basic single engine weighs approximately 150 pounds dry and 170 pounds at burnout. It has a cylindrical space envelope of 3 feet 6 inches diameter by 5 feet 9 inches length. The complete propulsion system (including the two engines and the tanks) weighs approximately 725 pounds dry, 19649 pounds wet, and 795 pounds at burnout. The preliminary design layout of the A-4 propulsion system is shown in figure 3-10.

Note that for the A-3 and A-4 engines a slightly smaller nozzle expansion area ratio has been specified than for the A-2. While all three upper stages operate in the vacuum and can use the largest practical expansion area ratio for best performance, other considerations will influence the ratio actually chosen.

System Operation

The propulsion system is designed to start automatically upon a signal from the guidance system. During main-stage operation, engine

Table 3-5.-7.5K A-4 Stage Engine Operating Parameters

[Vacuum conditions]
Engine (pressurized gas-feed and throttlable):		Main valve pressure drop	psi	4
		Calibration orifice pressure drop.	psi	8
Thrust	7500	Mixture ratio control reserve	psi	10
Nominal total multiple-firing		Oxidizer tank pressure	psia	165
duration at full thrust	410	Total oxidizer weight ( 410 sec full
Specific impulse	320	thrust duration for 2 engines, plus
Oxidizer $\mathrm{N}_{2} \mathrm{O}_{4}$ :		0.8 percent residual)	lb	10560
Density	90.88	Oxidizer tank volume (including
Flow rate	12.78	2.5 percent ullage volume).	$\mathrm{ft}^{3}$	120
Fuel $\mathrm{N}_{2} \mathrm{H}_{4}$ :		Nominal pressurant (helium) flow
Density	63.25	rate (assuming tank ullage
Flow rate	10.65	temperature $700^{\circ} \mathrm{R}$ ).	lb/sec	0.0225
Mixture ratio	1.2	Total pressurant weight (assuming storage bottle final temperature
Thrust chamber (ablatively cooled and radiation cooled on nozzle extension):		$191^{\circ} \mathrm{R}$, pressure 400 psia , plus 2 percent reserve)	lb	12.95
Thrust lb 7500		Pressurant storage tank:
		Volume	$\mathrm{ft}^{3}$	4.3
Specific impulse	320	Pressure	psia	4500
Injector end pressure	110	Temperature	${ }^{\circ} \mathrm{R}$	560 max.
Nozzle stagnation pressure	100
Oxidizer flow	12.78	Fuel side (pressurized by heated helium):
Fuel flow	10.65
Mixture ratio	1.2	Injector pressure drop	psi	25
c* efficiency	98	Inlet manifold pressure drop	psi	4
$c^{*}$	5540	Line pressure drop.	psi	4
$C_{f}$ efficiency	Percent	Main valve pressure drop	psi	4
$C_{f}$	1.858	Calibration orifice pressure drop	psi	8
Contraction ratio	2	Fuel tank pressure.	psi	155
Expansion ratio	35	Total fuel weight ( 410 sec full
Throat area, $A_{t}$.	40.4	thrust duration for 2 engines.
L*	32		lb	8840
Nozzle contour	70 percent bell	Fuel tank volume (including 2.5 ullage volume).	$\mathrm{ft}^{3}$	143.5
Thrust vector control:		Nominal pressurant (helium) flow rate (assuming tank ullage temperature $700^{\circ} \mathrm{R}$ ).
Minimum acceleration	$\mathrm{rad} / \mathrm{sec}^{2}$		lb sec	0.025
Maximum velocity	deg/sec	Total pressurant weight (assuming
Displacement
deg.	$\pm 7$	$191^{\circ} \mathrm{R}$, pressure 400 psia , plus
Oxidizer side (pressurized by heated helium):			1b	14.4
		Pressurant storage tank:
Injector pressure drop	25	Volume	$\mathrm{ft}^{3}$	4.77
Oxidizer dome pressure drop	3	Pressure	psia	4500
Line pressure drop	5	Temperature	${ }^{\circ} \mathrm{R}$	560 max.

thrust level and mixture ratio are controlled continuously through the engine control package by the guidance and propellant utilization systems. Upon a shutdown signal, engine shutdown is effected. The propulsion system is capable of restart an indefinite number of times. It can be operated at any thrust level between 10 percent and full thrust. Figure 3-11 shows the operational sequence of the A-4 stage engine.

Figure 3-9.-A-4 stage engine and propulsion system schematic diagram.

Figure 3-10.-15K A-4 stage propulsion system preliminary design layout.

Figure 3-11.-A-4 stage engine operational sequence.

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