from Public Record Office file DSIR 23/30814. Crown Copyright.

INTRODUCTION

Liquid hydrogen is amongst the most energetic chemical fuels which can be used. for rocket propulsion. In devising a research programme to investigate such an application It becomes evident that solutions to many of the problems involved can only be satisfactorily proven by the construction, launching and testing in flight of a vehicle employing liquid hydrogen as a fuel. Typical of such problems is the satisfactory arrangement of the tank structure and insulation to cope simultaneously with the effects of aerodynamic heating and the very low boiling temperature of liquid hydrogen, without incurring excessive weight penalties either in structure and insulation or in "boil-off" of liquid hydrogen. Another example is the problem of metering propellants and control of engine thrust and mixture ratio under conditions of acceleration and temperature changes in the propellants.

It was therefore decided to investigate the suitability of the ballistic research vehicle "Black Knight" as a first stage to carry a liquid hydrogen fuelled upper stage for research into these problems. As this work proceeded, it became evident that the development of a liquid hydrogen upper stage conferred on Black Knight the capability to launch earth satellites. It therefore became of interest to see how far this capability could be extended. by developing Black Knight into a heavier first stage, with a greater take--off thrust, by fairly simple modifications.

This note describes these possibilities and the sizes and characteristics of the upper stages considered. Estimates have been made of the stage weights and performances, leading to payload weights in a 300 nautical mile polar orbit. It has been assumed that the first stage developments (Section 2) would. be limited by a maximum tank diameter of 4 ft 6 in., fixed by the existing launching pad installation and that the propulsion unit would have to comprise components or multiples of the present Gamma 301 engine, using H.T.P. and kerosene. In the second stage (Section 4.1), the propellants are taken to be hydrogen and. oxygen, although fluorine is a somewhat more effective oxidant. For example, its use would increase the payload (in a 300 n.m.orbit) of the best vehicle described in this note (Scheme 4(a)) by about 20%. However, this would involve the solution of many problems of compatibility and toxicity and it was decided that the gain In performance was not sufficient to offset these disadvantages. Propulsion systems considered. for the second stage included pump and pressurised tank feed of propellants and either one or four combustion chambers and nozzles.

The trajectory calculations (Section 3) involved the use of a period of coasting up to apogee when either the second stage would be restarted or a third stage would be used to provide the final velocity increment. The solid propellant third stage was assumed to have characteristics within the range of performance of present solid propellant rockets. The guidance equipment would be carried in the second stage since, if carried in the third stage it would imply means of steering the stage, not a welcome complication in the case of a solid propellant rocket. Moreover, the payload. would be directly reduced by the guidance equipment weight.

It has been attempted throughout to make conservative estimates, so that the payloads quoted represent minimum values. In the tables of results, derivatives are given for the variation of payload with stage characteristics so that the payload estimates maybe modified as better estimates of these characteristics become available.

So far, these estimates have been based mainly on weight and performance assessments. Many aspects of the designs remain to be investigated in detail, including the aerodynamic, control and structural problems, the questions of impact range and guidance accuracies and the ground handling problems. It is hoped that the conservative approach adopted. here will accommodate any increases in weight which may occur during the solution of these problems. Such limited study as has been made (not reported here) has not indicated any serious likelihood of the estimates being grossly in error.

2 CONSIDERATIONS OF VEHICLE SIZE AND DEVELOPMENT

In developing a satellite launching vehicle, it is important to have the first stage available for flight tests before the upper stages have been completely developed. Tests may then be made using dummy upper stages of correct stiffness and mass distribution. These will show whether the first stage flight is satisfactory, before flight tests of fully operating upper stages are attempted. Black Knight was chosen as the basis of the study reported in this note because it is a developed, very reliable and relatively inexpensive vehicle. The increases in size and thrust of Black Knight which could be contemplated were therefore limited strictly to what could be achieved using either existing components (in the case of the engine) or existing equipment and manufacturing techniques (i.e., launchers, servicing towers, welding apparatus, etc.).

During the initial study of Blue Streak1 in the role of a satellite launching vehicle, an enlarged version of Black Knight was proposed as the second stage. In that study the considerations enumerated above suggested a maximum possible tank diameter of 4. ft 6 in. and this has been adhered to in the present study.

Since the development of the gamma 301 engine for Black Knight (thrust range 17,000 to 21 ,600 lb) the engine manufacturers have put forward proposals for future development work. These include raising the thrust of the 300 series engines to 25,000 lb and the development of a new turbo-pump unit to supply six chambers instead of four. Development firings on the present chambers, to demonstrate the feasibility of increased thrust, have been conducted at thrust levels up to 1.25 times the normal maximum thrust. Thus with a six chamber arrangement, a thrust level of approximately 40,000 lb at taketoff appears to be within reach at the cost of developing a completely new turbo-pump unit. A larger increase of thrust over that of the standard Black Knight could be achieved by arranging the components of two complete 301 engines in a propulsion bay of. the new diameter (4 ft 6 ins.). This engine, tentatively called the Gamma 401, would have two turbo-pump units and eight chambers identical to those to be used in Black Knight at a thrust level of 25,000 ib, giving a total take-off thrust of 50,000 lb.

This level of thrust probably represents a reasonable upper limit of development without either a major redesign of components or the use of different, larger components. In passing, it may be noted that this thrust level may be practically doubled by the use of four Stentor large chambers and turbo-pumps, a development which would be comparable with that of the Gamma 301, which was also developed from Stentor components. Such a level of thrust would, however, imply a larger tank diameter than 4 ft 6 in., thus precluding the use of much well-tried ground equipment. Nevertheless, the same principles could well be used in handling a vehicle of this size and its consideration should not be overlooked if the vehicles described. in this note do not provide capabilities compatible with such requirements as may arise.

For the purposes of the present assessment, therefore, vehicles with four levels of take-off thrust have been considered. To provide a basis of comparison, Version 1 is the standard Black Knight, with a tank diameter of 3 ft and a sea-level thrust of 21,600 lb. Versions 2, 3 and 4 have a tank diameter of 4ft 6in. and sea-level thrust of 25,000, 40,000 and 50,000 lb respectively. The take--off weight has been derived assuming a thrust/mass ratio of 1.25 and the propellant and upper stage weights have been derived by the methods described in the following paragraphs. Sketches of the vehicles, with leading dimensions, are given in Fig.1.

3. DETERMINATION OF TRAJECTORY AND STAGE MASS-RATIO REQUIREMENTS

Before considering detail design aspects of the various vehicles, it was necessary to establish the characteristics of a reasonably efficient launch trajectory, in order to fix suitable values of the ratio of mass at the beginning to that at the end of each period of thrust and the thrust/mass ratio of each stage. This had to be done bearing in mind certain assumed major limitations of the vehicle design such as length/diameter ratio and engine performance.

A circular polar orbit at 300 nautical miles altitude was considered, since this is commonly used. in launcher vehicle assessment work and. facilitates comparison with other designs. To achieve this orbit, the trajectory was assumed to consist of two consecutive periods of thrust arranged. so that the following coast flight would carry the vehicle to an apogee at the desired orbital altitude. A third. period of thrust would then follow, in which the vehicle would be accelerated to orbital velocity. Since the first stage performance would be largely limited by the known characteristics of Black Knight, the main variables affecting the trajectory were the "turn-over" programme adopted and the weight and thrust of the second stage. The thrust level of the third stage was not of significance since it was only used at orbital altitude to give, theoretically, an impulsive addition of velocity. In practice, the burning time will be finite but the errors in the orbit parameters are small on this account anti have been ignored.

Further details of the method of trajectory calculation are given in Appendix I and Figs.2 and 3 show some of the results, which are given more fully in Table 1. Further results in respect of trajectory profile and impact ranges are discussed briefly in Appendix 3. From the trajectory calculations, It was decided, as discussed in Appendix 1, to base the rest of the assessment on vehicles for which the ratio of total mass above first stage to launch mass is 0.174 and having a second stage thrust of twice the total mass above the first stage. For these vehicles, the attitude assumed at the end of the first stage "turn--over" programme is 40 to the launch horizontal. The trajectory calculations also yield a value of 2.87 for the ratio (mass at start/mass at end) of the second thrust period and values of .273 or 1.88 for the same ratio for the third thrust period, depending on whether this is provided by a solid propellant third stage of specific impulse 250 lbf sec/lbm or by relight of the second stage, which is assumed to have a specific impulse of 400 lbf sec/lbm.

In order to apply these results to actual designs in which the masses the empty stages might vary and to allow for possible differences in first stage engine performance, the following procedure was adopted. The basic vehicle chosen from the trajectory calculations had, for the first stage, a value of 1/(0.174 + 0.081) = 3.93 for the ratio () of mass at start to mass at end of burning (i.e. launch weight/(total upper stage weight + empty first stage weight)). Also, the sea--level specific impulse (I) was taken to be 214 lbf sec/ibm. The stage velocity increment was assumed to remain constant if the ideal velocity increment, , remained. unchanged. This enabled a new value of ? to be derived if the specific impulse were changed. For each launch mass specified, the mass at end. of first stage burning could then be calculated, the difference between these masses being the first stage propellant capacity. Knowing this, a weight could be estimated for the empty first stage, as described in Appendix 2, and this was deducted from the total weight at the end of first stage burning, so giving the second stage start mass. Use of the mass ratio of 2.87 (derived from the trajectory calculations) for the second stage gave the propellant consumed in the same way as for the first stage, hence enabling a weight estimate to be made for the empty second stage. This weight was subtracted from the total at the end of second stage burning to give the weight at the start of the third stage for which the pro-cess was repeated, using the mass ratio of 2.73, to derive a payload mass. A similar process was used for the two--stage vehicles and in each case it was assumed that the payload fairings would be jettisoned at the start of the coasting period.

The major characteristics of the vehicles derived by the above procedure are given in Tables 2 and 3 using the weights estimated for the different vehicles, as described in the next Section and in Appendix 2.

4. VEHICLE DESIGN AND WEIGHT ESTIMATES

Having decided the possible first stage developments as outlined in Section 2 and, on upper stage weights and. thrust/mass ratio 'by the procedure described in Section 3, the next step was to define the second stage configuration. It was considered that a comparison should be made between turbo--pump and pressurised tank systems of propellant feed, since, superficially at least, the latter appears simpler to develop, at the expense, however, of a possible weight penalty on account of the higher tank pressures which would be required. Further, a single chamber would make better use of the available cross--sectional area of the vehicle. Thus, for the same expansion ratio (and specific impulse), the throat area would be larger and the chamber pressure to give the required thrust would be lower, This would reduce the tank pressures required, compared with the four--chambered version. A comparison was therefore included between four chambers and a single chamber in the case of the pressure feed system.

4..1 Second stage propulsion systems

Three variants of the second stage engine and propellant feed systems were examined:-

(a) an engine having four chambers with turbo-pump feed of the propellants,

(b) an engine having four chambers with pressurised tank feed, and

(c) a single chambered engine with pressurised tank feed.

It was assumed that a combustion efficiency of 97% of theoretical, based on frozen equilibrium values, would be achieved. Thus, the theoretical specific impulse would need. to be 410 lbf sec/lbm to ensure the practical value of 400 used in the trajectory calculations. To obtain this theoretical specific impulse, the thrust coefficient for the nozzle would need to be approximately 1.76 at the proposed mixture ratio of 5:1, oxygen to hydrogen, by weight. An approximate value of the nozzle area ratio could also be obtained, depending on the chamber pressure, which remained to be derived. The maximum possible area at the nozzle exits was determined, allowing clearance for coolant manifolds, propulsion bay structure and angular movement of the chambers through 10 for control purposes, the nozzles being assumed pivoted at the throat. This led to values of the nozzle throat area for the two variants having pressure feed, (b) and (a), and hence to minimum values for the chamber pressures required to achieve the thrust levels associated with each upper stage mass. To derive the tank pressure, an arbitrary allowance of 25 p.s.i. was added. to the chamber pressures estimated in this way, to allow for pressure drops in the propellant feed lines, coolant passages and injectors.

In the case of the turbo--pump variant, (a), the chamber pressure was assumed to be 300 p.s.i.a., which appears to be about the maximum which can be used in association with a "no--loss" or "topping" turbine system of pump drive. The much higher chamber pressure of this variant implies a smaller throat area and, assuming the same value of thrust coefficient (1.76), a smaller exit area than could have been accommodated within the vehicle diameter. Nevertheless, the same practical specific impulse was adhered to (400), in order to simplify the comparison. It should be remembered, however, that advantage may be taken of the extra area ratio available to increase the performance of the stage if required. and this is referred to again later, in Section 5. The main characteristics of the three propulsion variants are given in Table 4.

In order to determine the length of the structure joining the first and second stages of the vehicle, it was necessary to estimate the lengths of chambers and nozzles. The chamber was assumed to have a value of L* (Volume/ throat area) of 25 ins, and a contraction area ratio of 225:1, the contraction being a conical frustum of half angle 15. The nozzle was also assumed to be a conical frustum of half angle 15. Total chamber plus nozzle lengths are given in Table 4 and it was assumed in the case of variant (a) that the chambers would be disposed around a single, centrally mounted, turbo-pump unit.

4.2 Weight estimates

Since detailed, accurate estimates are beyond the scope of the present work, the attempt has been made to take rather conservative values, In the hope of avoiding undue optimism in the magnitude of the predicted payloads. Details of the assumptions made in estimating the weights are given in Appendix 2. Weight breakdowns of the four versions of the first stage and the three variants of the second stage are given in Tables 5 and 6 for the three stage and two stage vehicles respectively. Broadly speaking, the weights of the first stage have been based on knowledge of the actual weight of Black Knight and its equipment. The second stage estimates have been largely based upon current information from the U.S.A., as far as possible using that which relates to methods of pressurisation, insulation etc. of which there is already some practical background of experience.

5 DISCUSSION OF THE VARIOUS DESIGNS

Examination of the payload weights derived in Tables 2 and 3 and summarised in Table 7 below discloses several major trends.

TABLE 7

Payloads in 300 nautical mile orbit

1st stage version Launch mass (lb) 2nd stage variant No. of chambers Propellant feed Payload (lb)
2 stage 3 stage
1 17,280 a
b
c
4
4
1
Pump
Pressure
Pressure
-ve
-ve
-ve
88
18
56
2 20,000 a
b
c
4
4
1
Pump
Pressure
Pressure
-ve
-ve
-ve
102
56
76
3 32,000 a
b
c
4
4
1
Pump
Pressure
Pressure
+30
-ve
-ve
324
169
248
4 40,000 a
b
c
4
4
1
Pump
Pressure
Pressure
80
-ve
-ve
377
187
289

(i) The variant (a) with a turbo-pump fed engine in the second stage shows a distinct advantage over the pressure-fed variants (b) and (c). Tables 5 and 6 show this to be due entirely to the extra weight of the tanks and pressurizing equipment of variants (b) and (c) compared with the relatively small weight of the turbo-pump in version (a).

(2) Of the pressure-fed variants, that with a single chamber and nozzle (c) gives a larger payload. This is because the area available for the nozzle exit is more efficiently used in this instance, which results in a lower chamber pressure being required, with consequently lower tank and pressurisation weights. It does not follow that a similar advantage over variant (a) would be shown by a single chamber fed by a turbo--pump. The chamber and tank pressures would not change in going from four chambers to one and the only significant difference would arise on account of the length of the bay between stages required to accommodate the engine length. This might well give the four chamber design an advantage over the single chamber design for turbo--pump engines.

(3) The ratio of payload mass to launch mass increases with launch mass. This result arises because of the fixed weights, in the second stage, of equipment and (at constant diameter) of the payload nose fairing. This result is well known but lends force to the argument for making the launch mass as large as possible.

(4) The three stage vehicles, as expected, give substantially larger payloads than do the two stage vehicles, despite the lower specific impulse of the solid propellant rocket providing the final velocity increment compared with the oxygen/hydrogen engine in the second stage. On the present analysis of two stage vehicles, only turbo--pump engines in the second stage are capable of ensuring some payload in a 300 nautical mile orbit, the pressure-fed variants having far too large a second stage empty mass, all of which has to be put into orbit.

A further consideration in the choice of type of vehicle lies in the length and length/diameter ratio of the vehicles. This has already been mentioned in Section 3 and Fig.1 gives dimensions of all versions with turbo-pump propellant feed in the second stage (variant (a)). It may be noted (Table 4) that the use of a single chamber in the second stage (variant (a)) adds appreciably to the length and. hence probably to the aerodynamic, control and structural problems. Fig. 1 also shows that the increase in launch mass from 32,000 to 40,0OO lb also gives a substantial increase in length and length/diameter ratio.

On the question of the absolute values of the payloads, it is fairly clear that only version 4, of launch mass 40,000 lb, gives reasonable assurance of having a satisfactory payload capability as a two stage vehicle (80 lb in a 300 nautical mile orbit) while offering the possibility of a quite substantial payload (377 lb) when a solid propellant third stage is added. The point may be made again that the weight estimates are thought to be conservative and some confirmation has recently been received that this may well be so. In the case of the first stage, a study3 by Messrs. Saunders-Roe of the design of a version of Black Knight with 51 in. diameter tanks suggests that the empty weight of version 2 derived in Table 5 might be reduced by 150 lb. For the second stage, studies by Saunders Roe4. and Rolls Royce5 of possible oxygen/hydrogen third stages for the Blue Streak satellite launching vehicle may be compared with second stage version (a) and again suggest that the empty weights quoted in Tables 5 and 6 for the second stage may be appreciably higher than those of a final design.

A further consideration which affects the payload is that of stage specific impulse. It is obvious that the first stage engine performance is relatively fixed. in this respect and values based on actual measurements have been quoted. In the case of the second stage, however, the specific impulse value used (400 lbf sec/lbm) is quite conservative and in the case of the turbo-pump variants, some increase may be foreseen, since the area ratios available are considerably more than have been used. For example, use of "bell" nozzles instead of 15 cones, but of the same length, could double the area ratio approximately and increase the S.I. by perhaps 4%. A further increase could. be obtained, at the expense of some separation bay structure length and weight, by expanding the nozzles up to the full area available, giving an area-ratio of approximately 4.0:1. Similarly, the third stage performance might well be increased by 5-10% by the use of somewhat more advanced techniques, though still well within present knowledge.

In order to facilitate the computation of the effect on payload weight of these and other changes, some values are given in Table 8 of the derivatives of payload weight with stage empty weight and specific impulse, assuming throughout that the stage ideal velocity increments remain unchanged. Consider scheme 4(a), the largest first stage combined with the turbo-pump fed. second stage. As designed, this would give a payload of 377 lb. If it is assumed that both first and second stage empty weights could be decreased by 5lb then an increment of about 30 lb would accrue. Similarly, if we assume increases of specific impulse in the second and third stages of 5%, a payload increment of 80 lb would result, giving a possible total increment of 110 lb and a total payload of about 490 lb. While it is felt that these improvements are well within the possibilities of development of the vehicle, it is considered that the first vehicles flown should not be expected to achieve much more than the 377 lb payload derived in Table 2.

To conclude this discussion, it should be remembered that many aspects of the designs remain to be investigated in detail, including the aerodynamic, control and structural problems, the questions of impact range and guidance accuracies and the ground handling problems. While preliminary study does not indicate that any of them are insuperable, many of them will involve a great deal of work and may well change many of the weights which have been estimated In this note. However, as discussed earlier, the total weight is not expected to exceed the values given.

6 CONCLUSION

This study of the possible use of enlarged versions of Black Knight together with a second stage using liquid oxygen and liquid hydrogen as propellants indicates that such a vehicle can place a payload of about 80 lb into a circular polar orbit at an altitude of 300 nautical miles. Addition of a third stage using a solid propellant increases the payload to about 375 lb.

The advantages of turbo-pump feed for the second stage engine compared to the use of pressurised tanks is quite clear at the stage sizes considered and the payloads just given are for the turbo--pump version. A pressure-fed version would not be capable even of placing the second stage in orbit and in the three stage version, would substantially reduce the payload.

The results of the trajectory calculations for this class of vehicles show that the second stage thrust should be at least twice the total initial mass of the upper stages and payload. This means that the engine developed for this second stage might well have applications for larger upper stages on other vehicles.

Estimates of the effects of change of stage empty weights and. specific impulses are given and are used to estimate the increments in payload which would accrue if less conservative stage characteristics were to be obtained. in practice, compared with those used here. This estimate indicates that a payload of around 500 lb should be attainable by the largest three stage vehicle considered, after some development and experience in the use of vehicles employing liquid hydrogen.

LIST OF REFERENCES

No. Author Title, etc
I Saunders-Roe Ltd. Black Prince, a launching vehicle for an earth satellite.
Saunders Roe Ltd. Publications Dept. T.P. No.435, May 1960.
2 Bristol Siddeley Engines Ltd.,
Power Division
Gamma. Proposals for continued development, 1963/64 - 1966/67. April 1962.
3 Westland Aircraft Ltd., Saunders-Roe Division Black Knight with increased diameter tanks.
Westland Aircraft Ltd. Report No. SP558, May 1962
4 Westland Aircraft Ltd., Saunders-Roe Division Blue Streak satellite launching vehicle. Design study of LOX/LH2 Third Stage.
Westland Aircraft Ltd. Report No. SP510, April 1962.
5 Rolls Royce Ltd.,
Advanced Projects (Rockets)
Liquid oxygen - liquid hydrogen rocket engine for a third stage of the "Blue Streak" satellite launcher. December 1961.
6 Robinson, H.G.R. The ballistic test vehicle "Black Knight". R.A.E. Tech. Note No. G.W.503. December 1958.

from Public Record Office file DSIR 23/30814. Crown Copyright

Dimesioned sketches of these four designs can be seen here.

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