Basic technical informations about A-4/V-2

DESCRIPTION OF THE GERMAN LONG RANGE ROCKET A-4

The long range rocket is a liquid fuel rocket designed to carry a payload of one ton a distaace of 300 kms. 
The rocket is made up of the following main components :
(a) The ballistic nose and warhead. 

(b) The control compartment.
(c) The centre section or fuel tank bay.
(d) The auxiliary motor unit and fuel pumps. 

(e) The rocket motor and venturi.
(f)  The tail unit and thrust ring.

 

THE BALLISTIC NOSE AND WARHEAD

 

The warhead is a sharply pointed cone. The side casing is of mild steel plate, which is braced internally by struts. At the rear end the casing is welded to a steel ring to which is also welded a circular end plate with a central hole for filling. About 3 ft. from the nose there is a hole leading to an internal pipe which runs through the warhead and the control compartment to the top of the alcohol tank for the purpose of pressurising this tank for the first 40 seconds of flight of the projectile.
Two diametrically opposite slinging points are provided in the side casing of the warhead and there are two more in the end plate itself. The warhead is attached to the control compartmeat by means of bolts and this junction forms one of the two transport joints of the rocket when on the Meilerwagen.
A central exploder tube, with electrical fuzes at its forward and rear ends, runs through the warhead. Each fuze has two inertia switches so positioned that they operate whatever the angle of impact. In addition, the nose fuze has a nose switch mounted on a steel tube which projects beyond the nose housing.

 

THE CONTROL COMPARTMENT 

 

The control compartment is situated immediately in rear of the warhead. It is a truncated cone, 4 ft. 7,5 ins. long, divided into four equal sections by radial plywood sheets. Relative to the plane of any two of the stabilising fins of the rocket, the plywood sheets are oriented at 45°.

Compartments 1 and 3 are each provided with two hinged panels held together by four fasteners. This permits easy access to the automatic pilot and alcohol tank inlet and the radio equipment, the hand operated stop cock for filling the alcohol tank pressurising bottles and the pressure gauge for reading pressure in the bottles. For insertion of the ground connector plugs in compartment 2, a spring-loaded hatch is provided in the main panel. There is also a  hatch for access to the main distribution box safety switch which is located in compartment 2.

Compartments contains:

1, Radio equipment for velocity measurement (Doppler) (Ortler or Verdoppler)

    Radio equipment for fuel cut-off signal (Honef or Kommandoempsfaenger)

    Emergency fuel cut-off equipment (NotBrennschluss)

    Two 16 V batteries (Bordbatterie)

2, Main junction box (Hauptverteiler) and distribution box safety switch (Bordautomat)

    Ground connector plug sockets (Stotz Stecker)

3, Fuze arming unit (Sterg)

    2 gyros, (Vertikant and Horizont ) on base plate ( Richtgeberplatte)

    Integrating accelerometer (I-Geraet)

    Time switch (Zeitschaltwerke)

    Control amplifier (Mischgeraet)
    50 V H.T. nickel iron battery (Kommandogeberbatterie)

    Alternator with regulator (Umformer mit Regler)

    Alcohol fuel tank inlet, gauge, stop cock

4, Receiver for radio control line (Leitstrahl)

    3 compressed air bottles (Zusatzbelunftung)

    2 alternators with regulators

    High pressure valve (Hochdruckventil)

 

CENTRE SECTION OR FUEL TANK BAY

 

The cenre section houses the alcohol and oxygen tanks, the forme being forward of the latter. The shell structure around the tanks is constructed in two halves and is made of steel reinforced by longitudinal and circumferential stringers.

The alcohol tank is constructed from light alloy sheet and is tapered towards the foward end of the rocket. The top of the tank is fitted with an inspection cover, a fuelling inlet, and an electrically operated pressure valve on the pipe running through the control compartment and the warhead. At the base of the tank is a pressure operated outlet valve and a drainage valve. The outlet valve leads via bellows to allow for expansion and contracting, to a delivery pipe which runs through the oxygen tank to the alcohol pump. Access to the alcohol drainage valve is by means of a hatchin the side of the shell between fins 2 and 3.

 The oxygen tank is also constructed from light alloy sheet. The tank is filled by means of a fuelling connection with a hand operated valve fitted to the tank outlet. An internal stack pipe which reaches to the top of the tank leads via a valve to a vent an the tail unit of the rockct. This pipe allows for venting of the tank during fuelling. A connection from the base of the stack pipe to the heat exchanger unit allows pressurisation by oxygen gas during flight. Access to the oxygen fuelling inlet is by means of a panel in the shell structure of the tail unit. The tanks and the alcohol pipe running through the oxygen tank are heavily lagged with glass wool.

 

THE AUXILIARY MOTOR UNIT AND FUEL PUMP

 

The purpose of the auxiliary motor unit is to provide steam for the turbine which drives the alcohol and oxygen pumps supplying fuel to the main combustion chamber of the rocket. The steam for the turbine is generated by the reaction of concentrated hydrogen peroxide and a catalyst, sodium permanganate. The fuels are stored in tanks which are filled through hatches in the main tail unit of the rocket . The hydrogen peroxide tank has a capacity of 126 litres. The sodium permanganate tank has a capacity of 11 litres. Provision is made for the drainage of these tanks, if, for any reason, fuel has to be removed from the rocket. Steam produced from the auxiliary combustion chamber drives the turbine to which are connected the alcohol and oxygen pumps.
Exhaust from the turbine is led via the heat exchanger into the rear of the rocket, where it escapes through cowlings in the tail unit. The heat exchanger vaporises some of the liquid oxygen, which is then used for pressurising the oxygen tank.
Oxygen is pumped from the oxygen tank through the oxygen main valve and subsequently by 18 pipes to 18 roses in the main cambustion chamber.
Alcohol is pumped via two pipes, each subsequently becoming three, to the base of the venturi, whence it passes between the walls of the venturi, for cooling purposes, to the jets in the combustion chamber.
The complete auxiliary motor unit, together with the turbine and pumps, is mounted on a braced steel framework bolted to the rear of the tank bay. This framework also supports seven compressed air bottles which pressurise the sodium permanganate and hydrogen pcroxide tanks and operate various valves which control the flow of fuels. An auxiliary electrical distribution box is fitted at the base of the oxygen tank.

 

THE ROCKET MOTOR AND VENTURI 

 

The combustion chamber and venturi constitute one welded steel assembly. The combustion chamber has 18 open-ended cups facing trarwards.
The oxygen is fed to the roses direct from the oxygen main valve. Alcohol, pumped to the base of the venturi, travels between the walls of the venturi to the alcohol main valve  which operates under pressure to allow alcohol to pass to the jets in the burners.
The venturi has rows of film coolant holes spaced throughout its length. These allow alcohol to pass to the inner surface of the vcnturi for additional cooling purposes.

 

THE TAIL UNIT AND THRUST RING

 

The tail unit is that part of the shell structure which encloses the rocket motor and auxiliary motor units. The tail unit does not transmit thrust loads; it acts as a fairing for the motor units and a support for the stabilising fins and the thrust ring at the base of the rocket. At its forward end the tail unit is supported by a circular rolled steel angle frame. This frame is bolted to the rear of the tank bay forming the second transport joint for the complete rocket.
The four stabilising fins are fixed at right angles to one another. Two are in the same plane as the plane of rotation of the turbine rotor ; they are numbered 1 and 3 and are aligned on to the target immediately prior to launching. Fins 2 and 4 are at right angles to the plane of rotation of the turbine rotor and are consequrntly at right angles to the plane of the trajectory when the rocket is in flight. At the rear end of each fin is an outer vane.
The tail unit has numerous hatches which are used for fuelling and give access to various valves, servo motors and potentiometers. At the base of the rocket, between fins 2 and 3 there is a 5-way coupling which forms the connection between the rocket and the valve box. The 5-way cupling is protected by a cowling. Between fins 1 and 2 and 3 and 4, respectively, are vents for the exhaust from the steam turbine. The cowling between fins 1 and 2 contains a valve which may be used for topping up the oxygen tank if there has been undue wastage since main fuelling was completed. In the base of fin 4 is a socket for the emergency fuel cut-off line, whilst fin 1 is fitted with a spring-loaded take-off switch (Abhebekontakt).
The thrust ring at the base of the tail unit is a cast light alloy channel ring which mounts the four supports and brackets for the carbon rudders . These carbon rudders, together with the four outer vanes on the fins, are used for controlling the rocket in flight up to the fuel cut-off point. The vanes and rudders are actuated by hydraulic servo motors, mounted inside the tail unit on the thrust ring. Two of the carbon rudders, together with their corresponding outer vanes (1 and 3) which are geared to them, control the rocket for line and counteract any tendency of the rocket to roll or yaw. Any such tendency is detected by the roll and yaw gyro which operates
the servo units through the control amplifier in the control compartment and the potentiometers on the thrust ring. The remaining two carbon rudders control the pitch (Program) of the rocket as directed by the pitch gyro. These latter rudders are not connected with their corresponding outer vanes (2 and 4) which assist control surfaces 1 and 3 correcting roll. 


PERFORMANCE OF THE GERMAN LONG RANGE ROCKET A-4


THE TRAJECTORY

 

The trajectory of the A-4 rocket may be divided into two parts :
(a) from launch until fuel cut-off,
(b) from fuel cut-off until fall of shot.
During the period before fuel cut-ff, the flight of the rocket is controlled; after fuel cut-off the rocket is uncontrolled.
When the rocket leaves the launching table, it first climbs vertically for 4 seconds. It is then made to pitch in the direction of the target in accordance with  a predetermined programme, which is put into effect by a time switch. The pitch programme lasts for approaitnately 43 seconds and at the end of this rime the rocket is at an inclination of 47° to the vertical. Thereafter, until fuel cut-off, which is approximatdy 65 seconds after launching, depending on the range to the target, the rocket is held at constant inclination and the trajectory is almost straight. After fuel cut-of, the rocket follows the normal parabolic trajectory of a free projectile attaining a maximum height of the order of 70 to 80 kms.
The initial acceleration of the rocket relative to the ground is 1g. The acceleration increases during burning as the weight of the rocket and the resistance of the atmosphere decrease, until it is approximately 5g at fuel cut-off. Air resistance, encountered as the projectile dascends into the atmosphere again, causes retardation. The maximum velocity attainable by the rocket is 1600 m per second, failing to a terminal velocity of 800 m per second.
The three axes of the rocket (pitch, roil and yaw) are controlled gyroscopically during burning. It is the purpose of these controls, together with the thrust of the rocket motor, to bring the rocket to a predetermined point in space, so that it has a predetermined orientation and velocity. When this point is reached, fuel supply to the rocket motor is shut off (Brennschluss} and control over the flight path is relinquished. Thereafter, the behaviour of the prolectile is similar to that of a normal shell, the velocity attained at fuel cut-off being comparable with muzzle velocity.
The trajectory up to fuel cut-off is assumed to be constant for all rockets. To fire at a given range, it is therefore only necessary to calculate the required velocity at fuel cut-off and to set the integrating accelerometer or sequence switch (or radio fuel cut-off devices) accordingly. The setting values can be read from range tables. The rocket, as used by the Germans, was fired entirely according to range tables.
Corrections are made to both line of fire and range to allow for the effect of the rotation of the earth. These corrections vary acoording to latitude, range of employment and direction of the line of fire.

 

RANGE CAPABILITIES OF A-4

With methyl alcohol and no fuel cut-off (i.e., all fuel is consumed), the mean range achieved by the rocket is 295 kms.  with a so-called 100 percent zone of +- 35 kms. Thus, although it is possible to reach a target of 330 kms, it is not in fact advisable to engage targets at ranges exceeding 260 kms which is the maximum effective operational range when methyl alcohol is employed. At ranges greater than 260 kms., an increasing proportion of rounds will exhaust their stocks of fuel before attaining the velocity required to take them to the target. The minimum operational raage is 80 kms., and this is governed by the limitations of the time switch which prohibits fuel cut-off before 45 seconds burning.

 

THE EFFECTIVENESS AND ACCURACY OF A-4

 

Rounds set up for firing may suffer one of three fates :

(a) They may fail to take-off due to some technical fault. 

(b) They may rise but fail to behave normally, also due to some technical failure.
(c) They may behave as expected and fall in the vicinity of the target.
During the bombardment of Antwerp by the Germans, from December 1944 to March 1945, the proportion of rounds falling into these three categories was 17 per cent., 18 per cent. and 65 per cent., respectively.
Of the 65 per cent, which behaved normally, the scatter of the fall of shot about the mean point of impact (MPI) varies according to the method of control employed. The following table indicates the accuracy achieved with the varying forms of control during operations against ANTWERP at a range of 200 kms.

TYPE OF CONTROL                                                          MEAN DEVIATION (KMS.)
    Range            Line                                                                Range                        Line

Mechanical    Mechanical                                                        3.4+-0.2                4.6+-0.3

Mechanical    Radio                                                                2.9+-0.3                0.4+-0.1

Radio            Radio                                                                6.7+-0.8                0.4+-0.1

 

 

Dimensions:
over-all length 14.3 m
warhead 2.285 m
control compartment 1.4 m
fuel-tank section 6.225 m
tail section 4.395 m
length of fins 3.935 m
diameter of body, center 1.65 m
diameter across fins 3.555 m

Weights:
warhead 975 kg  (amatol 750 kg)
control compartment 480 kg
fuel tank section (empty) 742 kg
rocket engine 931 kg
tail section with fins 855 kg
fuel 8800 kg
ethylalkohol 75% 3900 kg
liquid oxygen 4900 kg
hydrogen peroxide 80% 175 kg = 131 litres 
calcium permanganate 27 % 16? kg = 11 litres
dry weight 4008 kg
take-off weight 12805 kg

Performance:

Turbine:
diameter of blades 47 cm
working pressure 21 atm
rot. 5000 rpm
performance 496 kW
spotřeba páry 1.68 kg/s

Oxygen pump:
impeller diameter 27 cm
rot. 5000 rpm
performance 239 kW
delivery 75 kg/s
delivery pressure 24 atm

 Alcohol pump:
impeller diameter 34 cm
rot 5000 rpm
performance 265 kW
delivery 50 kg/s
delivery pressure 25 atm

Motor:
length of combustion chamber 1.725 m
diameter, combustion chamber 0.94 m
diameter, throat 0.405 m
diameter, nozzle exit 0.735 m
fuel consumption 125 kg/s
burnig time 68 s
exhaust velocity 2000 m/s
temperature in chamber 2000 °C
pressure in chamber 14.5 atm
thrust, sea level 27 t  tj. 250 kN
thrust, 40 km up 32 t  tj. 315 kN
acceleration 1 --- 6 g
velocity, max 1700 m/s
operational range 290-306 km
maximum observed range 354 km
operational peak altitude 97 km
vertical peak altitude 180 km